Staff Instruction (SI)

Approval Procedures for Modifications and Repairs to Damage Tolerant Aircraft Structures

File No.: 5009-32-2 SI No.: 513-001
RDIMS No.: 654638-V2 Issue No.: 02
Issuing Branch: Aircraft Certification Effective Date: 2004-02-16

1.0 Introduction

1.1 Purpose
1.2 Guidance Applicability
1.3 Description of Changes
1.4 Termination

2.0 References

2.1 Reference Documents

3.0 Background

4.0 Approval Procedures - General

5.0 Approval Procedures - Static Strength

6.0 Approval Procedures - Damage Tolerance Assessment (DTA) and Airworthiness Limitations

6.1 General
6.2 Final DTA and Airworthiness Limitations Complete
6.3 Final DTA Incomplete
6.4 Exceptions with Respect to Initial Approval based on Preliminary DTA
6.5 Temporary Repairs

7.0 Commercial Aircraft Under an Approved Maintenance Program

8.0 Modification and repair data provided by the original aircraft manufacturer

9.0 Delegation Aspects

10.0 Headquarters Contact

Appendix A - "Simplified Methodology" - Damage Tolerance Assessment Guidelines for Cut-outs in Semi-Monocoque Fuselage Structure

1.0 Introduction

1.1 Purpose

This SI provides procedural guidelines concerning the approval of modifications and repairs intended for damage tolerant aircraft.

1.2 Guidance Applicability

This document is applicable to Headquarters and the Regional Aircraft Certification personnel, including delegates.

1.3 Description of Changes

This issue of SI 513-001 provides a correction to the example calculation presented in Appendix A.5.2 of this SI at Issue 01, but is otherwise unchanged. Issue 01 was published to replace Aircraft Certification Staff Instruction (ACSI) No. 14, Issue 1, dated May 4, 2001. The contents reflect the current practices concerning the approval of modification and repairs intended for damage tolerant aircraft.

1.4 Termination

This document does not have a terminating action. It will however, be reviewed periodically for suitability of content.

2.0 References

2.1 Reference Documents

It is intended that the following reference materials be used in conjunction with this document:

  1. Chapter 505 of the Airworthiness Manual (AWM) - Delegation of Authority;
  2. Chapter 511 of the AWM - Approval of the Type Design of an Aeronautical Product;
  3. Chapter 525 of the AWM - Transport Category Aeroplanes;
  4. Airworthiness Manual Advisory (AMA) No. 505C/1 - Design Approval Representative (DAR);
  5. Aircraft Certification Staff Instruction (ACSI) No. 11 - Approval of Airworthiness Limitations - Regional Aircraft Certification Projects;
  6. Aircraft Certification Staff Instruction (ACSI) No. 22 - Approval Procedures - Domestic Design Changes; and
  7. U.S. Department of Transportation, Federal Aviation Administration (FAA) Advisory Circular (AC):
    1. 25.1529-1 - Instructions for Continued Airworthiness of Structural Repairs on Transport Airplanes; and
    2. 43.13-1B - Acceptable Methods, Techniques, and Practices - Aircraft Inspection and Repair.

3.0 Background

Following the incorporation of any modification or for Damage Tolerance (DT) needs to be evaluated for its continued DT capability. DT requirements can be incorporated into the basis of certification of an aircraft in one of two ways:

  1. Transport Category aircraft certified to Chapter 525 of the Airworthiness Manual (AWM) (FAA 14 CFR Part 25 including amendment 25-45), or equivalent requirements, are required to demonstrate DT as part of their basis for certification. Some other aircraft categories have similar provisions; or
  2. Transport Category aircraft certified to earlier standards have in some cases had mandatory Supplemental Inspection Programs applied through Airworthiness Directives (ADs). These aircraft are considered damage tolerant for design details to which the Supplemental Inspections apply and for the modifications and repairs themselves where the modified or repaired configuration is inferior from a fatigue perspective to the original configuration. The discrete source damage requirements of subsection 525.571(e) of the AWM are not applicable.

Airworthiness Manual Advisory (AMA) No. 505C/1 identifies DT as a function for which the finding of compliance is not normally delegated to either a Design Approval Representative (DAR) or a Designated Engineer (DE). Conventional modifications and conventional repairs applied within a single frame/skin bay to semi-monocoque fuselage constructions, far removed from significant stress gradients, may be delegated to qualified individuals. These qualified individuals must have demonstrated a working knowledge of Damage Tolerance Assessment (DTA) and the analysis and approval must be undertaken in accordance with a methodology agreed to and accepted by Transport Canada Civil Aviation (TCCA), Aircraft Certification (refer to Section 9.0).

The Airworthiness Limitations necessary to support a damage tolerance design require approval by the Chief, Engineering in Headquarters (AARDD). The Regional Manager Aircraft Certification may also approve the Airworthiness Limitations necessary to support a Damage Tolerance submission for modifications or repairs to existing type certified designs, per the provisions in ACSI No. 11. The Regional Manager Aircraft Certification approval of an Airworthiness Limitation must be based on the acceptability recommendation of a TCCA DTA specialist (Headquarters or Regional), or, his or her own finding if he or she is a DT specialist.

Changes to a Supplemental Inspection Program require approval by the Chief, Continuing Airworthiness in Headquarters (AARDG).

Delegation of Safe-Life and DT compliance findings pertaining to section 52X.571 of the AWM to individuals and organizations requires endorsement from the Manager, Aircraft Structures in Headquarters.

4.0 Approval Procedures - General

The introduction of a modification or a repair to a damage tolerant aircraft needs to be assessed to determine whether it affects an existing, or introduces a new, Principal Structural Element (PSE). A PSE is defined as any aircraft part or design detail area of a large part such as the fuselage carrying flight, ground, or pressurization loads, whose failure could result in the loss of the aircraft. Where the modification or repair affects or has an impact on a PSE, a DT evaluation must be completed in addition to demonstrating that the changed design continues to comply with all the other requirements of the aircraft basis of certification. Where there is no PSE impact, it is only necessary to demonstrate that the changed design continues to comply with all the other requirements for the aircraft basis of certification in the usual manner.

Regional Aircraft Certification Engineers continue to be responsible for the issuance of Supplemental Type Certificates (STCs) and Repair Design Certificates (RDCs) according to the procedures in ACSI No. 22. The DT aspects of such approvals are subject to the following being met:

  1. Considerations identified in this document;
  2. Requirements for the approval of airworthiness limitations specified in ACSI No. 11; and
  3. Factors identified in ACSI No. 22 to assist in making the assessment of where a project is to be managed.

The approval of any structural modification or repair to a damage tolerant aircraft must include the following:

  1. A static strength assessment;
  2. new PSEs;
  3. Review and approval of the DTA by a TCCA DT specialist with specialization in DT Methods, or, approval by a person or organization delegated to make findings of compliance with DT requirements; and
  4. TCCA approval of new or revised Airworthiness Limitations, if found necessary.

Appendix A of this SI provides a simplified DTA methodology for demonstrating compliance with section 523.571 and 525.571 of the AWM. This simplified methodology may be applicable to small penetrations and repairs to pressurized semi-monocoque fuselage structure and should only be used where the applicability defined in Appendix A is completely satisfied.

5.0 Approval Procedures - Static Strength

Before an aircraft is returned to service, the static strength of any structural modification or repair must be completed and approved by either TCCA or an appropriate delegate.

6.0 Approval Procedures - Damage Tolerance Assessment (DTA) and Airworthiness Limitations

6.1 General

A DTA of a modification or repair may result in new or revised Airworthiness Limitation(s), or both, associated with the modification or repair. For cases where the DTA indicates that no new or revised limitations are required, the supporting DTA or rational must be included in the substantiation data for the approval. A DTA of the affected structure must normally be completed by the applicant and approved by TCCA, or a person or organization delegated DT signatory authority, before returning the aircraft to service (refer also to section 6.2). TCCA may allow for Airworthiness Limitations based on preliminary DT assessments, which are interim in nature pending completion of the final DTA and associated Airworthiness Limitations (refer to section 6.3). Note that regardless of whether an Airworthiness Limitation is intended as interim, it constitutes an Airworthiness Limitation. Allowing a time limited Airworthiness Limitation to expire without appropriate resolution normally results in the aircraft no longer meeting its basis of certification. Airworthiness Limitations arising from DTAs are normally specified in terms of Threshold and Repeat Inspection Intervals with prescribed inspection parameters including but not limited to: the type of inspection method to be utilized and a description of the inspection location. The prescribed inspection instructions should be sufficiently comprehensive as to enhance the probability of detecting any fracture. The inspections may be specified based on flight cycles, special mission cycles, landings, calendar time or other appropriate parameters identified as most limiting in the DTA. Airworthiness Limitations, required in support of modifications or repairs to DT aircraft, must be approved as revisions or supplements to approved Airworthiness Limitations documents applicable to the aircraft being modified. The revisions or supplements to approved Airworthiness Limitations documents must be referenced on the approval document for the modification or repair.

A serious airworthiness concern may exist if an applicant does not regularly finalize the necessary DTA within the appropriate interim period, or, requires recurring extensions to the interim period. In such cases, it may not be appropriate for TCCA to allow an applicant to continue to use the provisions of sections 6.3, 6.4 or 6.5.

6.2 Final DTA and Airworthiness Limitations Complete

In this case, all DTA substantiation is complete and required Airworthiness Limitations are in place before the aircraft returns to service. The DT analysis is submitted to TCCA for approval with the recommendation of a delegate, where appropriate, unless such approval is delegated to the person responsible for the repair design in accordance with procedures agreed with TCCA. The aircraft is returned to service when the analysis is approved, any necessary Airworthiness Limitation(s) have been approved and the modification or repair is approved. Some foreign regulatory authorities may require that their own registered aircraft have permanent Airworthiness Limitations in place before accepting a modification or repair made outside that country.

6.3 Final DTA Incomplete

In this case, TCCA would allow usage of Airworthiness Limitations that are interim in nature pending completion of final DTA(s) associated with the modification(s) or repair(s). The Airworthiness Limitations to be used in the interim must be submitted for approval by the applicant before the aircraft being returned to service. An interim Airworthiness Limitation should not exceed a 12-month calendar period. A preliminary DTA is required to support the validity of the Airworthiness Limitation in the interim. A qualified individual must perform the preliminary DTA to establish valid interim Airworthiness Limitations. The accuracy of the preliminary DTA required will depend on the criticality of the modification or repair. In all cases, the conservatism of the Airworthiness Limitation must be immediately obvious or be otherwise supported by analytical or test data and submitted to TCCA for approval, or a person or organization delegated damage tolerance signatory authority.

In the case of modifications and small repairs involving fuselage penetrations to fuselage semi-monocoque structure, restoration of static strength while ensuring that significant fatigue degradation has not been introduced would normally allow for a 6 to 12 month Airworthiness Limitation to be supported. For more critical cases, a preliminary analysis (supported by test if necessary) would be required before return to service of the modified aircraft; such a DTA will normally provide for revised inspection intervals and possibly revised inspections methods. For less critical cases, the DTA may show that existing inspection intervals and methods are still valid, or that no new inspections/methods are required.

Airworthiness Limitations must be specified and issued in accordance with procedures agreed with TCCA, or in an Airworthiness Limitation document that requires TCCA approval, and is referenced on the approval document. The Airworthiness Limitation Approval must include provisions that will ensure that a permanent Airworthiness Limitation will supersede the Airworthiness Limitation intended for the interim usage. This is normally accomplished by specifying:

  1. A chronological limitation on the validity of the Airworthiness Limitation used in the interim after which the aircraft is to be considered as no longer airworthy; and
  2. That the Airworthiness Limitation used in the interim must be superseded by an approved revision log (to the approved supplement or revision to the Airworthiness Section of the Maintenance Document housing the Airworthiness Limitations) specifying instruction for its removal or replacement.

TCCA may allow a 3-month extension to be applied to the Airworthiness Limitations used in the interim in exceptional circumstances if warranted. The extension would be approved and implemented in the same manner as the original Airworthiness Limitation used in the interim.

6.4 Exceptions with Respect to Initial Approval based on Preliminary DTA

An interim life limit may not be required if a methodology delineating the scope of the modification or repair has been approved together with an approved tracking system. The tracking system must be sufficiently comprehensive to ensure that there is confidence that the required Airworthiness Limitations will be in place within an appropriate time from the introduction of the modification or repair. The applicant must be able to show that he has the ability to complete the necessary analysis, the ability to generate any necessary Airworthiness Limitations documents and a reliable system for recording the outstanding analyses, as well as tracking and completing outstanding damage tolerance activities. The applicant should provide regular updates on the status of outstanding DT analyses. This approach would normally be limited to aircraft manufacturers having a well-established DTA working relationship with TCCA.

The aircraft is returned to service before the completion of the final supporting DT analysis on the basis of a static approval of the modification or repair and a preliminary DTA. The applicant must have shown that he can complete the DT activities within the period, not to exceed one year, for which he has established that no Airworthiness Limitation Inspections are required. The approval document must clearly identify the period during which the DT analysis must be completed. This case would not normally be applied to non-Fail-Safe feature PSEs unless agreed to by TCCA.

6.5 Temporary Repairs

Although the provisions of sections 6.3 and 6.4 for an interim life on a permanent repair could be used for a temporary repair for up to one year (if appropriate), temporary repairs are often associated with short-term solutions. Provisions to address temporary repairs should be included in a Delegates Engineering Procedures Manual or Design Approval Procedures Manual, as applicable. At the end of the interim life a temporary repair must either have been fully substantiated and approved as a permanent repair in the normal manner (completed final DTA and Airworthiness Limitations) or be removed from the aircraft.

7.0 Commercial Aircraft Under an Approved Maintenance Program

Commercial air operators may exercise control over the application of mandatory structural inspections through TCCA approved maintenance programs. The maintenance program may stand alone, or may include mandatory inspections or Airworthiness Limitation documents. There is a need to ensure that adequate control of mandatory inspections is continued when an aircraft is moved from one maintenance program to another, such as following a transfer of ownership. Pending the establishment of appropriate procedures for record keeping of each repair, care should be taken to ensure that all required inspections associated with modified or repaired damage tolerant aircraft are clearly identified as mandatory inspections.

One acceptable approach would be for an operator to accumulate required Airworthiness Limitations arising from modifications and repairs, and then to submit them for TCCA approval on a quarterly basis. Alternatively, the operator may elect to issue its equivalent approval documents under its delegation authorized by TCCA and in accordance with procedures agreed with TCCA.

8.0 Modification and Repair Data Provided by the Original Aircraft Manufacturer

Some Canadian manufacturers have prepared documents and associated data tracking protocols for Structural Deviations and Repairs. An example of this is the Structural Deviation, Inspection and Repair (SDIR) program applied to many Bombardier Aircraft. The SDIR facilitates tracking and provides a repository for data used in mandating type certificate holder aircraft specific production deviations, aircraft specific type certificate holder repairs, and third party repairs to a specific aircraft. TCCA may delegate certain structural deviations and repairs to type certificate holders based on approved methodology documents (also providing scope) and approved Structural Repair Manuals.

9.0 Delegation Aspects

TCCA does not normally allow the exercise of full delegation of findings of compliance related to DT and Safe-Life. For small modifications and small repairs to in-service aircraft, TCCA Aircraft Certification may allow exceptions. Such exceptions would only be applicable to conventional small modifications and conventional repairs to semi-monocoque fuselage constructions far removed from significant stress gradients. Typically, modifications such as antenna penetrations and repairs (within 1 fuselage bay) to holes, scratches and dents may be considered as small modifications or small repairs.

TCCA may grant delegation for such provided the following considerations are satisfied:

  1. Delegation is limited to an individual or individuals within an organization who have demonstrated expertise in DT Analysis/Safe-Life and compliance procedures;
  2. The delegate or delegated organization has prepared a DTA methodology document that has been approved by TCCA Aircraft Certification, Structures. The methodology shall ensure that conservatism is maintained and contains as a minimum:
    1. A detailed description of the scope of DTA/Fatigue to be performed;
    2. Detailed description of the methods to be used (derivation of loads, any simplifying assumptions, initial crack sizes, geometry factors used, critical crack length determination, special factors applicable to calculation of threshold and repeat inspection);
    3. Material data property sets and corresponding crack growth models and tools; and
    4. Aircraft applicability, aircraft zonal applicability, and restricted zones.
  1. The delegation authorization letter must reference the individual(s) that are making the finding of compliance to the DTA/Fatigue. The delegation letter must also reference the methodology document(s) that will be used by the authorized individuals; and
  2. All authorizations are subject to regular TCCA audits.

10.0 Headquarters Contact

For more information please contact:

Policy Standards Coordinator (AARDH/P)
Phone: (613) 990-3923
Facsimile: (613) 996-9178

Original signed by Gilles Morin for

Maher Khouzam
Chief, Regulatory Standards
Aircraft Certification Branch

Appendix A - "Simplified Methodology" - Damage Tolerance Assessment Guidelines for Cut-outs in Semi-Monocoque Fuselage Structure

A.1.0 Introduction

The simplified DTA guidelines provide an approach or methodology for demonstrating compliance with section 523.571 and 525.571 of the AWM. This simplified methodology may be applicable to small penetrations and repairs to pressurized semi-monocoque fuselage structure and should only be used where the applicability defined in this appendix is completely satisfied. Generally, small antenna penetrations that use standard industry design practices, materials and fasteners to restore structural integrity and are far removed from large cut-outs, or, significant stress and strain gradients, will qualify.

Two examples are:

  1. Example 1 described in A.5.1 below, uses the simplified methodology; and
  2. Example 2 described in A.5.2 below, investigates cracking along a doubler rivet line and is provided to demonstrate that results obtained are less restrictive than those obtained using the "simplified methodology" of example 1.

If the design were unconventional or otherwise ineligible for treatment with the simplified methodology to obtain conservative inspection intervals, a standard DTA assessment would be required. A standard DTA assessment of patch repairs and modifications to semi-monocoque fuselage designs will require consideration of doublers and detailed loads; cracking at the first line of rivets tends to be the limiting case.

A.2.0 Applicability

Considerations and parameters that must be satisfied before using the simplified methodology are:

  1. Agreement must be reached with the regional aircraft certification engineer or Headquarters structures engineer, or both, that usage of the simplified methodology is appropriate for the intended:
    1. Class of repair or modification; and
    2. Application location on the aircraft.
  1. Usage of the simplified methodology is limited to fuselage:
    1. Semi-monocoque construction affording redundant load path features at locations where the modifications or repairs are far removed from large stress concentrations such as doors, windows, wing frames and pressure bulkheads;
    2. Semi-monocoque construction at locations where the modification or repair does not degrade critical design details such as lap joints, circumferential fuselage joints and other significant discontinuities; and
    3. Constructions affording large critical crack lengths (greater than 1 bay). Aircraft designs that satisfy the 2 bay centre frame crack scenario generally afford such critical crack lengths.
  1. The simplified methodology is only applicable where a fuselage:
    1. Penetration is compensated using a doubler of equal gauge or greater which affords a double row of rivets satisfying standard design practices (refer to AC 43.13-1B); and
    2. Repair results in a cut-out or geometrical restoration not exceeding a 3" diameter circle. The restoration or cut-out must be compensated using a doubler of equal gauge or greater which affords a double row of rivets satisfying standard design practices (refer to AC 43.13-1B).
  1. The simplified methodology may not be used with "novel" or irregular design practices. An example of such a practice would be where the design allows for imparting excessive out-of-plane loads on the aircraft skin, or, inserting fasteners at non-conventional locations.
A.3.0 Simplified Methodology Description

The validity of this methodology is based on satisfying the following conditions and assumptions:

  1. The repair or modification analysis:
    1. Ignores doublers used as part of the repair or modification;
    2. Uses recognized and conservative material parameters for the crack growth model in the most critical in-plane material direction;
    3. Assumes diametrically opposed through thickness 0.025" cracks on both sides of the penetration or equivalent cut-out coincident with the geometrical restoration;
    4. Does not consider crack growth retardation effects;
    5. Does not consider bulging effects;
    6. Does not use averaged frame/skin bay stresses such as "Flugge" stresses; and
    7. Considers crack growth in the longitudinal direction or principle stress directions only.
  1. For calculating repeat inspection intervals, the analysis considers the geometrical crack growth path, which results in the shortest inspection interval coincident with the inspection method used.
  2. For the purpose of calculating threshold and repeat inspection intervals, the minimum visual inspectable crack shall be considered as 1.0" long for symmetrical crack propagation from a hole. If no hole is present, the minimum visual inspectable crack length to be assumed is 2.0".
  3. The overall design, with the exception of the penetration or cut-out itself, maintains fatigue stress concentration factors values comparable to adjacent type certificate holder design details.
  4. If crack growth is assumed along a rivet line, the existence of the rivet line is ignored.
  5. Threshold and Repeat Inspection Intervals are calculated using the following factors:

    Threshold Inspection = (afinal - ainitial)/(K1*K3*K4)
    Repeat Inspection = (afinal - adetectable)/(K2*K3*K4)
    Where      K1 = 2
                    K2 = 2
                    K3 = 1.0 low humidity environment
                          = 1.5 medium humidity environment
                          = 2.0 high humidity environment
                    K4 = Special Scatter Factor which accounts for 
                    unknowns (usage of this Special Factor must be 
                    agreed with TCCA
  6. The prescribed threshold and repeat inspections never exceed half (½) of the aircraft life regardless of the results obtained from the simplified analysis.
A.4.0 Simplified DTA Methodology Caveats

The following caveats apply:

  1. The design and quality assurance must ensure that the design is conventional in all respects;
  2. Ensure minimum edge distances, minimum fastener spacing and countersink depths are respected;
  3. Ensure compatibility of metals in contact (i.e. Galvanic Corrosion);
  4. Provide for bonding of doubler if possible;
  5. Ensure the design is capable of reacting to out-of-plane loads to frames or stringers without very significant flexure of the skin;
  6. Ensure rivet pattern is conventional and satisfies ultimate strength requirements;
  7. Incorporate a doubler design which "picks-up" on existing stringer/frame rivets in lieu of stopping short;
  8. Ensure ease of report legibility by: providing clear geometry, loads derivation, listing of assumptions, program input and output data, crack length versus cycle count plot, Beta versus crack length plot if possible, details of mandated inspections and special instructions;
  9. Investigate analysis sensitivities by performing parametric variations;
  10. Ensure correctness and conservatism of all of inputs including material properties data; and
  11. Provide a rationalization of the correctness of the results obtained.
A.5.0 Examples
A.5.1 Example 1 - Usage of the Simplified Methodology

Aircraft Type: Bombardier CL-600

Modification Description:

  1. VHF blade type antenna installation.
  2. 2" Feed-through at mid bay in the vicinity of fuselage station (FS) 454 between stringer 1 and 2.
  3. Full bay (frame to frame and stringer to stringer) internal doubler of 2024-T3XX.
  4. Doubler attached via double row of AD5 rivets respecting minimum edge distances and fastener spacing and ensuring "knife-edge" countersink depths are precluded.
  5. Transverse stringer-to-stringer stiffeners are provided to react overturning moments on VHF antenna to avoid excessive out-of-plane loads on doubler/skin.
  6. Antenna fasteners attach to 2024-T3XX "pads" or transverse stringer-to-stringer stiffeners to avoid larger stress concentrations.
  7. Antenna planform longest direction = 7.0".


  1. One solution is to use the Air Force Crack Growth Analysis Fracture Mechanics tool, called AFGROW, which is available for download from the Wright Patterson Air Force Base, Vehicles Directorate, Air Force Research Laboratory, United States Air Force web site at Note: AFGROW Version 4.0008.12.11, 6/16/03 is the reference for the following examples.
  2. Ensure material properties used are identical or conservative with respect to repaired structure:
    1. Select material data tab.
    2. Select Nasgro Equation option and OK.
    3. Select Read... tab.
    4. Select Open Material Database tab.
    5. Select appropriate Material, in this case 2024-T3 Clad Plate and Sheet L-T (conservative).
    6. Select OK top left of active window.
    7. Select Apply and OK.
  1. Input Stress Spectrum. AFGROW allows for input of a "spectrum file" or the selection of constant amplitude loading. For the CL-600, use of constant amplitude pressurization (Pr/t) is conservative:
    1. Select Constant Amplitude Loading.
    2. In the Stress Multiplication Factor field, input the "Pr/t" stress in KSI (11.8 KSI for this example).
    3. In the following window select block size = 10, select OK. This block size will improve the fidelity of the reported ½ critical crack size C.
  1. Input the model geometry. AFGROW provides a list of predefined geometries. User defined geometries must satisfy:
    1. Select Classic Models tab.
    2. Select double through crack at a hole (beta solution is shown as application defined).
    3. Select Dimension tab.
    4. In the Width field, input 100 inches, which assumes infinite plate width.
    5. In the Thickness field, input the fuselage skin thickness in inches ignoring the doubler thickness and pad-up thickness. In this example t = 0.045".
    6. In the Hole Diameter field, input equivalent size diameter of the hole encompassing the cut-out, in this example D = 2.0".
    7. In the Crack Length field, input the initial crack length. In this example the initial crack length is a = 0.025". For geometry with a single crack tip, this value specified would be 0.05".
    8. Select Load tab.
    9. Under Stress Ratio in the tension portion, enter 1.
    10. Select Apply followed by OK tabs.

Note that the reported ½ crack size C at failure is 9.65" with a corresponding pressure cycle count of about 22,000. AFGROW may also report a ½ crack size based on Kmax criteria; the lowest critical ½ crack length should be selected. The reported critical crack length is therefore equal to 2*C + Dcut-out = 21.3" If we had chosen the T-L direction as material properties in lieu of L-T, the critical crack length would have been 2*C + Dcut-out = 15" with a pressure cycle count of 40,000. Inspection intervals based on the more critical grain direction are required.

Since the skin crack is visually inspectable externally, the calculated threshold is deduced as 10,900 cycles, assuming for this example that K1=2, K3=1, K4=1. However, if the type certificate holder life of the aircraft is 15,000 cycles, the simplified methodology would require that the prescribed threshold be stipulated at 7,500 cycles.

The repeat inspection interval would be deduced in a similar manner with an initial crack length corresponding to the detectable crack length. The detectable crack size will depend on the nature of the non-destructive inspection (NDI) and associated instructions. It must be decided whether the inspection instructions will specify internal or external visual or other NDI as well as other options such as removal of the antenna.

Since the doubler in this case is internal, an internal skin inspection is undesirable as the doubler effectively hides the skin. Other factors such as the removal of the interior may also make this choice less desirable; had we considered the doubler in the analysis, we would have no choice but to inspect for the integrity of the doubler.

An external visual check will be investigated. A visual threshold of detectability for directed - location specific inspections is often assumed as 1.0" to 2.0". The maximum length of the crack that can exist under the VHF antenna allowing for detection is therefore = 2" + 7" = 9". Recall that 7" is the longest direction of the antenna planform.

Re-running the AFGROW analysis with a starting crack length = adetectable = 9.0" requires that the Crack Length field input be changed from a = 0.025" to a = 3.5". (2" hole + 2 * 3.5" = 9").

The result reported is < 1200 cycles to failure.

Repeat Inspection = (afinal - adetectable)/(K2*K3*K4)
= 1200/2
= 600 cycles

This result is probably much too restrictive to be practicable.

If instead, the mandatory inspections required removal of the antenna with a NDI inspection such as dye-penetrant, the result would have been a repeat inspection equivalent to the threshold inspection.

A.5.2 Example 2

This example uses the AFGROW "multiple load case" feature to calculate the crack growth from a rivet along the first row.


  1. Analysis of crack growth at the first row of rivets requires that the doubler load transfer be investigated.
  2. Investigation of a crack emanating from a cut-out (example 1) without consideration of doublers generally leads to more limiting inspection intervals than this example. Investigation of a crack emanating from the first line of rivets (per this example) should be performed when the design is unconventional.

For simplicity, we use the same model as above but with the following changes:

  1. Select Classic Models tab.
  2. Select SingleThrough Crack at a hole (beta solution is shown as application defined).
  3. Select Dimension tab.
  4. In the Width field, input the rivet pitch, say 4D, in this case if we use a 5/32" rivet, W = 0.625".
  5. In the Thickness field, input the fuselage skin thickness in inches. In this example t = 0.045".
  6. In the Hole Diameter field, input D = 0.156".
  7. In the Crack Length field, input the initial crack length. In this example the initial crack length is a = 0.05".
  8. Select Apply tab.
  9. Select Load tab.

At this point we must deduce the, bypass stress, reference stress, bearing stress ratio and tension stress ratios. (The bending stress ratio in the absence of flexural loads = 0.0). 

For this example, it is assumed that within the double row of rivets the doubler has reduced the far field stress of 11.8 ksi by 50%. We also assume that the load sharing between the first row of rivets and second row of rivets is 60%/40% (60% for the 1st row of rivets). The load transferred by the 1st row of 5/32" rivets in bearing is therefore:

Pin Load = 11.8 ksi * 0.625" * 0.045" * 0.5 * 0.6 = 100 lb.

Bypass Stress = the bypass stress about the 1st rivet row: 11.8 ksi - (11.8 ksi * 0.6 * 0.5) = 8.3 ksi.

Reference Stress = 8.3 KSI + 100 lb/((.625" * 0.045")*1000) = 8.3 KSI + 3.6 KSI = 11.9 ksi

Bearing Stress Ratio = 100 lb/ (1000*(0.156")(0.045"))/11.9 = 1.20

Tension Stress Ratio is calculated as 8.3/11.9 = 0.70

Note that the pin diameter was not used in the development of the reference stress, but was used in the bearing ratio calculation (refer to the AFGROW help file).

  1. Under Stress Ratio in the tension portion, enter 0.70
  2. Under Stress Ratio in the bearing portion, enter 1.22
  3. Select OK tabs.

Select the Spectrum icon:

  1. Select Constant Amplitude Loading.
  2. In the Stress Multiplication Factor field, input the reference stress as calculated above (11.9 ksi in this example).
  3. In the following window with block size > 10, select OK.

The result reported by AFGROW is 52,000 cycles, which is the number of cycles required to propagate the crack to the adjacent rivet hole. At this point, the rogue crack needs to reinitiate on the opposite side of the hole. Treating the effected rivet holes and crack as a single crack and repeating the AFGROW run (centre crack model with crack = 2C = (0.156" + 0.156" + 0.625") = 0.937" in 100" wide panel ignoring holes and doubler) yields an additional 40,000 cycle count before critical crack size is achieved. The total crack growth life is therefore 92,000 cycles. This result is conservative as it neglects the cycle counts required to re-initiate cracks at every fastener hole.

Had we selected a "double through crack" at a rivet hole, with C= 0.0025", the result obtained would have been 91,000 cycles, a measure of the durability or Multi Site Damage due to manufacturing flaws.


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