Damage Tolerance

History of Fatigue Failures and Evolution of Regulations

The safe life criteria was followed in the Aerospace industry before 1954 where the safety was ensured by planned retirement.

The fatigue design gained increased importance after the series of failures of the pressurized cabin aircrafts. The crash of two Comet 1 aircrafts in 1954 lead to airworthiness authorities mandating Failsafe design. However, there was no operational limit on the failsafe design certified aircrafts. Below are the details of the incidents

• 10 Jan 1954 : De Havilland D.H.106 ‘Comet 1‘ disintegrated mid-flight at an altitude of 30.000 ft and crashed into the Mediterranean sea off Elba with the loss of all the passengers and crew members.
• 8 Apr 1954: Another ‘Comet 1’ crashed into Mediterranean near Naples. Subsequent investigation revealed that the accidents were caused by failure of pressurized passenger cabin brought about by fatigue.

During mid-70s, the crash of US Air F-111 and the Boeing 707-300 Freighter lead to the mandatory Damage Tolerance based Inspection on the Aircraft structures

• 1976 : US Airforce F-111 fighter-bomber crashed while attempting a 4.0 g steady maneuver around 100 flight cycles.
• 14 May 1977 : Boeing 707-300 Freighter of Dan-Air Services crashed on approach to Lusaka International Airport in Zambia. The accident investigation revealed that the right hand horizontal stabilizer (horizontal tailfin) and the elevator had separated, causing the aircraft to pitch rapidly nose down and dive into the ground.

The Damage Tolerance Requirement (Safety by Inspection) was mandated by the regulatory authorities on both existing and future aircrafts. This lead to

• Amendment 25-45 to part 25 (on the new aircrafts)
• Advisory Circular 25.571-1 (on the new aircrafts)
• FAA AC 91-56- 12/1981 (on the existing aircrafts implementation by airworthiness directives (ADs)
• Supplemental inspection documents (SIDs) for specific airplanes of concern were developed by the manufacturers based on AC 91-56

In the late 80s and early 90s widespread fatigue damage and aging aircrafts safety gained attention following the incident of Aloha Flight 243 (Boeing 737). The aircraft suffered an midair explosive decompression resulting in the loss of upper fuselage during its flight from Honolulu to Hawaii. The accident was attributed to the presence of multiple cracks in the longitudinal Joints.

Following the Aloha Flight incident, the Widespread Fatigue Damage and Aging Aircrafts safety received the attention and it lead to the introduction of the following rules.

• Aging Airplane Safety Act 1991
• Aging airplane structure rulemaking
• Supplemental inspection program, revision to certain supplemental inspection documents: AD-mandated program
• Mandatory modification program: AD-mandated program
• Repair assessment program: Operational rule
• Corrosion prevention and control program: AD-mandated program
• Aging airplane safety rule: Operational rule and part 26 rule.
• Widespread fatigue damage (WFD) rulemaking.
• Amended §25.571 (Amendment 25-96)
• Introduced the term “WFD” into the regulations.
• Introduced damage tolerance certification requirement to show freedom from WFD up to the design service goal by full-scale fatigue test evidence

Most recent WFD rulemaking includes.

• Amendment to §25.571 (Amendment 25-132)
• Amendment 25-132 establishes Limit of Validity (LoV) for future airplane models.
• Include LoV in ALS
• Perform full-scale fatigue testing to validate the LoV relative to WFD

Damage Tolerance Philosophy

Damage Tolerance is defined by FAA as ‘The attribute of the structure that permits it to retain its required residual strength for a period of use after the structure has sustained a given level of fatigue, corrosion or accidental or discrete source damage’.

The period of time (FC, FH or both) established at design and/or certification during which the airplane structure is reasonably free from significant cracking is called ‘Design Service Goal’

Key elements of the Damage Tolerance include

• Residual Strength
• Crack Propagation
• Inspection Interval

The following terminologies (as defined by FAA or EASA) are widely used for the Damage Tolerance categorisation of structures

Principal structural element (PSE) — An element that contributes significantly to the carrying of flight, ground or pressurization loads and whose integrity is essential in maintaining the overall structural integrity of the airplane.

• Principal structural elements include all structure susceptible to fatigue cracking, which could contribute to a catastrophic failure.

Structural Significant Item (SSI) – SSI is defined as a principal structural element that could fail and consequently reduce the structural integrity of the airplane.

Fatigue Critical Structure (FCS) – is defined as airplane structure that is susceptible to fatigue cracking which could contribute to a catastrophic failure.

Primary Structure – The primary structure is that structure which carries flight, ground or pressure loads.

Secondary Structure – The Secondary structures are those structures which carries only air or inertial loads generated on or within the secondary structure.

Crack Propagation

Crack Propagation analysis is carried out to estimate the number of cycles (Flight cycles) required for the crack to propagate from initial crack length to the critical crack length

Initial flaw length is recommended to consider manufacturing defects. 0.05 inch or 1.27 mm is the standard initial flaw length practiced in the Aerospace industry.

Critical crack length is determined using the residual strength calculations

The crack growth rate da/dN under the fatigue loading is a function of stress intensity range  ΔK and the stress ratio R

The crack propagation lives are estimated using one of the following different power laws

• Paris Power Law
• Forman and Modified Forman Equations
• Elber Equation

Paris Power Law

Paris Power law is the most simplest form of power law to estimate the crack propagation life. It is expressed as

da/dN=C(ΔK)n
Where, 
The coefficient C and the exponent n are constants for a given material 
and cyclic stress ratio, R

The fatigue spectrum generally consists of cycles of different stress ratio. Therefore, usage of Paris law requires different coefficients C and n for different R ratios.

Further, the Paris law considers only the Region II of the sigmoidal curve. The Forman law overcomes this drawback by extending the region of applicability.

Forman Equation

The Forman equation accounts the behaviour in the region I and region III with slight modification to the Paris equation. The Forman equation is given by

da/dN=C(ΔK)n/[(1-R)Kc-ΔK]
Where,
	Kc is the critical stress intensity factor of the material

The modified Forman equation accounts the threshold behavior in the region I of the sigmoidal curve. The modified Forman equation is given by

da/dN=C(ΔK-ΔKth)n/[(1-R)Kc-ΔK]
Where,
 Kc is the critical stress intensity factor of the material

Residual Strength

The structural elements are designed for the ultimate loading but the daily loads or operating loads experienced by the aircrafts are much below the limit strength.

The strength of the component reduces in the presence of a damage (crack).

The residual strength of the structure should not fall below the regulatory requirement

The regulatory requirement for the residual strength is generally limit strength but depends on the structure.

Objective of the Damage Tolerance based inspection program is to detect the crack before it becomes critical and restore the structure to its original capability with the help of repairs.

Figure shows the reduction in the residual strength of the structure in the presence of the crack. The crack propagation curve is superimposed to show the reduction in the residual strength capability of the structure. The critical crack length corresponds to the point where the residual strength falls below the regulatory requirement. This is the point where the fast fracture is assumed to occur.

The repair before residual strength becomes critical restores the structure to its original capability.

Residual Strength

The critical damage length should be such that the residual strength remain above the regulatory requirements. The regulatory requirement requires that that the structures are capable of withstanding ultimate loads. Therefore, the aircraft structures are designed at ultimate loads. The daily operating loads (fatigue loads) experienced by the commercial aircrafts are much below the limit loads. With the propagation of a fatigue damage, the residual strength (remaining strength) reduces due to the reduction in the cross sectional area as can be seen in the below figure. The residual strength is not allowed to fall below the regulatory requirement. Therefore, it is necessary to have a maintenance program where the damage can be detected before it becomes critical. The regulatory requirement for most of the structures is the limit load. The pressurised structures have to meet 1.10 (FAR) or 1.15 (CS/JAR) times the differential cabin pressure depending on the  certification authority.

The critical damage (Crack) length is determined to meet the residual strength requirements using the following approach

• K_IC exceedance (for plain strain scenario)
• Crack Resistance Curves (R-Curves)
• Net Section Yield (Static yield)
• Feddersen Criteria

K_IC exceedance is mostly used to calculate the residual strengths of the thick parts and corner flaws. The corner flaws behave more or less like a plain strain condition. Therefore, most industries use K_IC exceedance to evaluate the residual strength of initial corner flaw.

R-Curves and the Feddersen criterial are mostly used for the plain stress components. Additionally, net section yield is also considered for the plain stress structures to ensure that the component is not separating statically.

Inspection

Damage detection by planned inspection ensures the safety of the Aircraft.

The damage in the Aircraft should be detected and repaired before it becomes critical.

There should be sufficient number of inspections directed at a structural element to detect the damage.

If the damage is not detected by one inspection then the subsequent inspection/s should ensure the damage is detected before it reaches the critical size.

The first inspection is called the ‘Threshold of Inspection’.

The interval between the subsequent inspections is ‘Repeat Inspection Interval (RII)’ or simply ‘Inspection Interval’.

The RII depends on the structural design, loading level and the method of inspections.

The method of inspection and the RII play a vital role in the Maintenance cost of the aircraft. The visual inspections are cost-effective in comparison to the NDT techniques. Therefore, visual inspections are always preferred. 

Inspection Methods

Methods of Inspection is chosen based on various aspects. The Non destructive Inspection (NDI) can detect small flaws. However, they are expensive. Visual Inspections are most preferred type of inspections due to economic reason.

The inspection techniques used for Aircraft maintenance are listed below.

Visual Techniques

• General Visual Inspection (GVI)
• Detailed Visual Inspection (DET)
• Bore scope Inspection

NDI Techniques

• Liquid Penetrant Inspection
• Eddy Current Inspection (HFEC, MFEC, LFEC)
• Ultrasonic Inspection
• Acoustic Emission Inspection
• Magnetic Particle Inspection
• Radiographic Inspection (x-Ray)

The Eddy Current Inspections can be High Frequency Eddy Current (HFEC), Mid Frequency Eddy Current (MFEC) or Low Frequency Eddy Current (LFEC). The LFEC is used mostly to detect the sub surface cracks (i.e. the crack in the part is completely hidden by another part and there is no possibility to inspect the crack from the opposite side). The HFEC is used to detect the surface cracks. Ultrasonic inspection is also used for subsurface inspection.

The x-Ray is not a preferred method of NDI for the components made of parts from different materials as it incorrectly show other part as a crack due to the frequency set to detect crack in one material does not work for other material.

Widespread Fatigue Damage (WFD)

ASTM defines Fatigue as a localised phenomena. However. the Fatigue can be global as in Widespread fatigue. The Widespread fatigue can be multiple site damage or multiple element damage.

The following definitions are obtained from the FAA 25.571

Widespread fatigue damage (WFD)—The simultaneous presence of cracks at multiple structural locations that are of sufficient size and density that the structure will no longer meet the residual strength requirements.

Multiple site damage (MSD)—A source of widespread fatigue damage characterized by the simultaneous presence of fatigue cracks in the same structural elements.

Multiple element damage (MED)—A source of widespread fatigue damage characterized by the simultaneous presence of fatigue cracks in similar adjacent structural elements.

Inspection start point (ISP)—The point in time when special inspections of the fleet are initiated because of a specific probability of having a multiple site damage/multiple element damage condition.

Structural modification point (SMP)—The point in time when a structural area must be modified to preclude WFD.

WFD (average behavior)—The point in time when without intervention. 50% of the fleet is expected to develop WFD for a particular structure.

Widespread Fatigue Damage

Widespread Fatigue Damage (Courtesy: FAR 25.571)

According to the latest FAA/CS 25.571 amdt 132, it mandatory to specify the Limit of Validity (LoV) in the Airworthiness Limitations Section (ALS2) for the new aircrafts.

The LoV is the period of time (in flight cycles, flight hours or both), up to which it has been demonstrated that WFD is unlikely to occur in an airplane’s structure by virtue of its inherent design characteristics and any required maintenance actions..

An airplane may not operate beyond the LoV unless an extended LoV is approved.