Crack assessment

Citation

Ford, T. (2004), "Crack assessment", Aircraft Engineering and Aerospace Technology, Vol. 76 No. 6. https://doi.org/10.1108/aeat.2004.12776faf.003

Crack assessment

Crack assessment

Following the events that occurred in the 1980s major differences became evident in the recognition and treatment of aircraft fatigue damage. Although concern had existed for some time it was the Aloha incident that focused attention on the problems faced, in particular, by ageing aircraft. Now, life extension programmes and new aircraft designs have to consider the consequences of widespread fatigue damage (WFD). The main issue of WFD in ageing aircraft is the occurrence of multiple damages at adjacent locations which influence each other. Two types of multiple damage are known; one is multiple site damage (MSD), which is characterised by the simultaneous presence of fatigue cracks in the same structural element and the other is multiple element damage (MED) in which the simultaneous presence of fatigue cracks occur in similar adjacent structural elements. Both MSD and MED are a source of WFD which is reached when the MSD or MED cracks are of sufficient size and density that the structure will no longer meet its damage tolerance requirements.

Early in these activities an interim solution was defined for 11 aircraft types in the category of ageing aircraft or those soon to reach their design service goal (DSG). The Airbus A300 is typical of the latter and life extension aspects became of interest to the company when early A300s reached more than 75 per cent of their DSG in the 1990s. The actions comprised the following: periodical review of the in-service experience regarding structural damage (service bulletins); introduction of a corrosion prevention and control programme (CPCP); assessment of the fatigue life of structural repairs; establishment of a supplemental structural inspection programme (SSIP) to reach the new safety standards; and assessment of the structure regarding WFD. Aspects of these activities are supported by the results of full scale fatigue tests (FSFT). The life extension programme for the A300 was completed in 2001, which evaluated as well as included FSFT results, which included specific component tests for areas susceptible to local and WFD, and the tear down investigation of a retired aircraft. The latter allows the correlation between analysis, test results and in-service results. Airbus has a history of FSFT on all its aircraft to support certification, which are performed as multifunction tests.

Realisation

One of the ageing aircraft issues relates to the then reviews which are handled by groups composed of the manufacturers, operators and airworthiness authorities for each of the types concerned. These are tasked to review all structurally related service actions/bulletins and determine which require mandatory terminating action or enforcement of special repetitive inspections. The original equipment manufacturers (OEM) should undertake this activity to select candidate service actions and the operators are requested to submit information on individuals fleet experience related to the candidate service actions/bulletins which should be followed by the final selection of recommendations of mandatory actions. This decision is based on potential airworthiness structural concerns, reliability of inspection, frequency of occurrence, and adjacent structural damage. This review is periodically repeated.

The introduction of a corrosion prevention and control programme (CPCP) has been dealt with a previous issue of “Aircraft Engineering and Aerospace Technology”. Another action is the continuous airworthiness of existing repairs which has been identified as a significant concern. The final draft report and subsequent rules for this appeared in the late 1990s, which requires the establishment of repair assessment guidelines for existing repairs. The A300 Repair Assessment Guideline (RAG) had been started by Airbus at this time, followed by the development of a PC tool which automates the complete assessment process. For final verification of the fatigue life and threshold determination in the RAG two large repairs were installed in the A330/A340 FSFT before starting fatigue testing. These were a longitudinal lap joint repair in the centre fuselage/ wing specimen and a skin repair in the rear fuselage specimen.

The SSIP has largely overtaken previously issued guidelines and can be considered a “live” document which has kept in touch with the development and maturity of the A300 based on in-service experience. The A300 SSIP is in two parts, the first relating to significant structural details subject to service bulletins and the second part to details subject to the Fleet Leader Programme. The SSIP breaks the structure down into a number of inspection zones. For each zone the inspection requirements applicable are detailed giving the mean fatigue life, threshold crack propagation information and the repeat interval.

Widespread fatigue damage

As indicated, a major concern is interaction between multiple fatigue cracks, whether of MSD or MED origin, which may compromise the residual strength of an aircraft before their detection at typical inspection intervals. Study has shown that in most cases, by the time that MSD has become in-service detectable it is already too late. Extremely small damage cannot be tolerated within the operational life of the aircraft if in fact it will cause the originally certified residual strength capability to degrade below the limit load regulatory requirement. The question has been raised in the past as to why tiny cracks at fastener holes should cause concern when the structure has been designed to sustain large damage. It has to be remembered that tiny cracks at fastener holes are really equivalent to these cracks plus the diameter of the fastener holes and therefore, small cracks can have substantial effect on the large damage capability.

A number of analytical methods have been developed to determine the effects of MSD on lead crack residual strength, one being based on the concept that link-up of the MSD cracks with the lead crack will occur when plastic zones from the two crack trips touch each other. This is based on net section yield failure between the crack tips for ductile materials. Plastic zone sizes can be related to crack tip intensity factor and these factors can be calculated in cracked stiffened structures. A reasonable assumption for ductile alloys such as 2024-T3 is that link-up of the lead crack with the MSD crack will occur when the intact ligament stress between the two crack tips reaches the typical yield strength of the material, leading to the intuitive criterion. This is one of the analysis tools that have been determined to assess the effects of MSD on lead crack residual strength.

A frequent query is how many MSD cracks does it take to substantially reduce the residual strength of the lead crack. With a second MSD crack the panel is assumed to be stable, but if the process is continued by adding MSD cracks it will become apparent that after considering four or five cracks of this nature, the residual strength does not recover after link-up of the first crack.

One example is the DC-10 fuselage which was designed for a large damage capability which was substantiated by finite element cracked panel analysis validated by test. The tests were performed on large panels which had already been subjected to at least three times initial design lifetimes of cycle loading, the aircraft being initially certified for limited crack arrest capability and damage state. The main concern here is how long can this crack arrest, large capability last before it becomes affected by the onset of WFD with continued operation of the aircraft beyond its original design life goal.

For the simple case of a one bay skin crack in a stiffened panel the residual strength will be given by the upper curve on the left. The peak of this residual strength curve, based on skin fracture criterion, is shown well above the limit load stress, and the crack arrest capability exists up to and above the limit load. If the aircraft now continues in operation beyond the initial design life goal and MSD begins to formulate, there may be a possibility that the peak of the residual strength curve will drop below the limit load as shown on the right. It appears that the loss in lead crack residual strength in the presence of MSD is very sensitive to structural geometry. In some cases substantial MSD can be tolerated and in other cases even the smallest MSD will substantially reduce residual strength capability. This means that all areas of the structure susceptible to MDSD should be evaluated for the size of MSD which would cause the certified residual strength to drop below the regulatory levels.

In consequence the size of MSD crack which would cause lead crack residual strength to be degraded below the regulatory levels needs to be calculated for those areas susceptible to this type of cracking. A crack growth evaluation needs to be made to determine the time to reach this point, which requires an estimate to be made on an initial flaw size which would be representative of the initial quality of the structure. Cracks of a size which can drastically reduce lead crack residual strength have a low probability of detection and therefore, the onset for WFD has to be established so the aircraft can be modified before another problem occurs.

Manufacturers progress

The maintenance of structural integrity encompasses developments in testing and in-service experience learned over the years. Applications of structural improvements incorporating the fatigue knowledge outlined and a wealth of other expertise are used by manufacturers to develop inspection, maintenance and repair procedures with the data also helping the enhancement of analytical methods for predicting behaviour.

Typical of this experience is the Boeing 777 on which flight cycling on the FSFT facility began at the beginning of 1995 and ended in March 1997. The 777 was tested to 120,000 cycles which is three times the design service objective (DSO) or 60 years. However, results of this test do not provide an alternative to inspections required by the maintenance programme and separate testing was done for the horizontal stabilizer and elevator, because the attachment to the body is determinate, and for the nose and main landing gears as these are safe life.

At the full scale test facility, repeated spectrum loading was applied with each test spectrum made up of a block of 5,000 flights with these being continually repeated. Each flight consisted of a ground-air-ground profile that was made up of ground handling, taxi-out, take-off, cabin pressurization, cruise, depressurization, descent, buffeting, landing, and taxi-in. In addition to verifying the aircraft to the DSO, the fatigue test verifies that the customer will not experience unusual repair costs associated with structural damage. Test results continue to be used in the design phase of 777 derivative aircraft to maintain structural efficiency.

Technology standards

Development of these standards as a means of analysing durability, fatigue and damage tolerant behaviour is an important element of the structural improvement process. Damage tolerant standards were developed in the 1970s and applied to subsequent aircraft, the corrosion prevention standard was released in the late 1980s and the design value standards at the end of the 1990s. The Boeing durability standards development has contributed to the structural improvement process and is shown by the fact that the number of design change required was reduced after test compared to those of earlier models (757,767), even though the 777 was cycled torn 50 per cent beyond the earlier aircraft. These tests have been instrumental in corrosion prevention, repair of composites, access and maintenance, collection of ageing fleet and durability data, and establishment of a database of customer inputs.

Particular areas of this development on the 777 include a lighter weight and more corrosion resistant titanium alloy on the engine pylon, and the aft upper spar fittings which are integral with the upper spar chords. On earlier designs, the mid spar fittings were mechanically fastened to the upper spar chords. The attachment itself was a potential source of crack initiation and corrosion and was difficult to inspect. Extensive use is also made of carbon fibre reinforced plastic (CFRP) on the empennage.

Specific attention has been paid to design of the floor beams with corrosion resistance a main consideration and CFRP used extensively. Also, the beams allow metal bolt-on repairs similar to those for the empennage. The fuselage of the 777 has incorporated experience from the ageing fleet of older models, including improved fatigue life, crack growth-resistant properties and toughness. Notably, there are thicker gauges of material used in lap splices, resulting in reduced stress levels and increased fatigue life. In some areas of body skin, 2524-T3 clad sheet is used which is tougher and has a lower crack growth rate than 2024 at the same stress intensity level. In the wing structure, high strength 7055-T7754 plate is used for the upper wing skin and 7155-T7751 extrusion for the upper spar chords, both high strength and corrosion-resistant materials developed during the 1990s. These are some examples of the way in which continuing test and service experience are improving design as well as looking ahead to the incorporation of materials which will exhibit even better performance.

Terry Ford

Further readingCEAS Forum. Life Extension – Aerospace Technology Opportunities. The Royal Aeronautical Society 1999.Maintaining Safety in the Millenium. The Royal Aeronautical Society 1997.