First part of report on RAeS/CEAS forum: life extension - aerospace technology opportunities

Citation

(1999), "First part of report on RAeS/CEAS forum: life extension - aerospace technology opportunities", Aircraft Engineering and Aerospace Technology, Vol. 71 No. 4. https://doi.org/10.1108/aeat.1999.12771dac.001

First part of report on RAeS/CEAS forum: life extension - aerospace technology opportunities

First part of report on RAeS/CEAS forum: life extension - aerospace technology opportunities

Keywords: Fatigue, Aircraft

This conference was jointly organized by the Royal Aeronautical Society and the Confederation of European Aerospace Societies (CEAS) and brought together experts involved in these technologies with the opportunities to disseminate information on these issues. It addressed interaction between the research community, civil and military aerospace manufacturers and operators, and maintenance certification and regulatory authorities. Organized in several sessions, the papers covered aspects of ageing aircraft, repair and life extension, monitoring and NDT, modelling and simulation, and fatigue life improvement.

A comprehensive and informative Keynote Address by a former Chief Scientist to the FAA set the scene for the conference. T. Swift began his presentation outlining the profound effects of the Aloha accident in 1988 and the subsequent formation of the Structural Audit Evaluation Task Group (SAETG) which developed guidelines to establish the onset for Widespread Fatigue Damage (WFD) in those areas of the airframe susceptible to Multi-Site-Damage (MSD) and Multi-Element-Damage (MED). In Figure 1 residual strength is shown as a function of flight time for single site local damage which may occur inadvertently during the service life. This is shown by the dotted line. The presence of MSD was thought to degrade the residual strength capability more rapidly, indicated by the solid line.

Despite every effort being made to ensure that fatigue cracking will not occur, due to inadvertent circumstances, it does happen and in-service inspections are expected to find damage before catastrophic failure. Each principal structural element (PSE) is analyzed under spectrum loading to develop a crack growth curve. Continued operation of aircraft beyond the service life goal exposes the airframe to the formation of WFD. In-service undetectable MSD can substantially reduce residual strength of the single lead crack used to establish the inspection programme. Loss in residual strength of the single lead crack can reduce the safe crack growth period and for this reason MSD cannot be tolerated within the operational life of the aircraft for structures where it can cause residual strength loss below the regulatory level.

After the Aloha accident the FAA Ageing Aircraft Research Programme was formulated. The most probable cause of this accident was linked up MSD in the fuselage skin at the upper row of countersunk rivets of the lap splice at stringer ten on the left side of the aircraft. Some research and development had been done before Aloha but it is firmly believed to be the wrong approach and current efforts by the FAA believe that MSD cannot be tolerated within the operational life of the aircraft, if in fact it will cause the original certified residual strength capability to degrade below the limit load regulatory requirement.

The real problem is that even the smallest MSD cracking can substantially reduce residual strength capability for which the aircraft was certified (Figure 2). Tiny cracks at fastener holes are really equivalent to these cracks plus the diameter of the fastener holes. Analytical methods have been developed to determine the effect of MSD on lead crack residual strength, and examples are given. It has become apparent that the loss in lead crack residual strength in the presence of MSD is very sensitive to structural geometry.

Figure 1 Residual strength capability and resulting inspection actions

Figure 2 Effect of MSD on lead crack residual strength

Examples of accidental damage and fatigue damage are given as well as flows. For the latter, time to the onset of WFD would be calculated by conducting a crack growth evaluation starting with multi-site equivalent quality flaw sizes. The time to the onset of WFD would be the time it takes the equivalent initial quality flaw size to reach a length that would cause the certified lead crack residual strength to be degraded below the regulatory level.

Ageing aircraft

The majority of the papers concerned with this aspect were from Airbus, one being an Aerospatiale Assessment Approach for WFD in Ageing Aircraft Structures. This gave details of the various aspects of the A300 life extension programme including maintenance, testing, susceptibility, etc. High time A300 aircraft are now reaching their Design Service Goal (DSG), i.e. the life for which the aircraft was initially certified. An Extended Service Goal (ESG) has been defined to substantiate at least 25 per cent additional life.

An important part of this is the Airbus Widespread Fatigue Damage (WFD) assessment programme. Various means are used including full-scale fatigue tests. The AS300 fatigue test was performed 20 years ago but the test specimen is still available and further detailed tear-down investigations will be performed and some possible WFD susceptible areas identified.

An overall prediction method has been developed in recent years, notably identifying the link between initiation and propagation of WFD. Also, life to initiation of a crack has a random aspect because of the scatter existing in many influential parameters. For overall assessment a simulation process built around the Finite Element Method has been developed.

The calculation method must be supplied with realistic input data and small coupon tests representative of the A300 structure are being performed. Large representative tests will also be used if the full scale fatigue test did not supply sufficient data.

Two examples of WFD assessment are first, the front pressure bulkhead where MSD could develop in the outer ring splice, at fastener rows and in other areas. This component is only subject to cabin pressure and the specimen experienced 135,000 simulated pressure cycles during the full scale fatigue test. No significant damage was found during test and subsequent tear-down but more detailed tear-down investigations will be done to check whether tiny cracks existed that were not detected earlier. The second example is the circumferential joint at frame 54 which is subject to high secondary bending, particularly in the butt strap. MSD could possibly occur at rivet holes and MED could also be met by failure of successive stringer couplings along the joint. Multiple damage was found during full scale fatigue test and led to immediate improved design of the joint. Three types of test will be performed for WFD evaluation: fatigue test on small coupons; fatigue test on flat stiffened panels and residual strength tests on curved panels.

From British Aerospace Airbus came a contribution also concerned with the A300, on opportunities and a proposed means of wing structure justification. In the early 1990s a general extended operations programme was expected to be initiated soon but it was also thought that a certain number of the B4 variant would reach their DSGs before the date of 2000. These activities include the Supplemental Structural Inspection Programme (SSIP) and regular reviews of existing maintenance programmes. The latter result in periodic updates of Service Bulletins, corrosion control programs, and the SSIP.

For an aircraft wing, locations potentially susceptible to MSD include chord wise splices; stringer run-outs at tank end ribs; and rib to skin attachments. A further location for MSD on the A300 wing is the top skin repair plate on some early versions of this aircraft.

Work on WFD has included full scale fatigue test with simulation of a longer period of operation, and a forthcoming teardown inspection. A research project is also underway. A number of large panel tests are being performed by BAe Airbus, simulating the top skin chord wise joint, bottom skin chord wise joint, and the stringer run-out features. One panel of each will be used in a fatigue test with a spectrum loading for the A300B-600 variant which will be the first to exceed the design life goal.

A more accurate estimation procedure of the fatigue life of a multiple fastened structure has been developed which will be used for MSD assessment. Thus, a procedure for examining MSD/MED behaviour based upon all possible crack configurations that arise naturally forms the basis of the BAe Airbus procedure rather than an approach which attempts to use "typical" crack patterns, with all the uncertainty that entails.

Lessons learned from recent European research projects on ageing aircraft were detailed by Peter Horst of the University of Braunschweig. Different research projects on ageing aircraft illustrate how the European approach varies from that of the USA. This paper is concerned with the former, probabilistic approach. The Monte Carlo simulation is used for the process initiation and growth of MSD scenarios.

Some possible reasons for the origin of MSD-like scenarios, mainly design mistakes, high loads, aggressive environments, bad manufacturing, and possibly the wrong material suggest that a probabilistic approach is preferable. A Monte Carlo simulation of the fatigue related MSD problem provides very good and interesting results for the assessment of MSD, although further development is required mainly with respect to small-scale effects and very small probabilities.

Details of the Airbus full scale fatigue tests were provided by Daimler Chrysler Aerospace Airbus. It notes that all Airbus types other than the A300 were certified after 1978, i.e. after FAR 25.571 Amendment 45 which requires justification of sufficient damage tolerance behaviour of the structure. All Airbus aircraft have full scale fatigue tests (FSFT) to support certification, which are performed as multi-section tests. A minimum of two life times was simulated for all FSFT. The long term full scale fatigue testing allows the detection of nearly all areas which may be fatigue sensitive within the design service goal, since the scatter in loads and material data is covered sufficiently. The development of the damages over the test progress is shown in (Figure 3). Justification of sufficient fatigue life for structural repairs, concessions and reworks is also possible.

The extensive results for FSFT allow significant conclusions for the structure and planning of life extension activities. For the A300, together with other data, extended service goals (ESG) have been defined. Different A300 variants have different ESGs, including such factors as airworthiness, economic aspects, and modification states.

Following Aloha and the ageing aircraft issue, an interim solution was defined for 11 aircraft types, which included the A300. The following were defined: periodical review of the in-service experience regarding structural damage (review of service bulletins); introduction of a corrosion prevention and control program (CPCP); assessment of the fatigue life of structural repairs; establishment of a supplemental structural inspection program (SSIP) to reach the safety standard according to FAR 25.571 Amendment 45; and assessment of the structure regarding widespread fatigue damage (WFD). The planning of the A330 life extension programme involved extensive testing and evaluation of areas, both the 18 potentially susceptible to WFD and other areas.

New FAR regulations a year ago contain the requirement for full-scale fatigue tests and tear-down inspections which are underlined by the Airbus philosophy of these tests up to approximately 2.5 life times or more which are performed for all types which allows to extend easily the original design service goal.

Structural requirements and repair considerations

The Defence Evaluation and Research Agency (DERA) at Farnborough described the life extension aspects of the structural design requirements for British military aircraft. The paper showed how the lessons learned have been carried forward. For military aircraft, changes in usage can have powerful influences on fatigue consumption.

Military structural airworthiness requirements for combat aircraft and rotorcraft use the safe life design philosophy, whereas military transport and light trainer types are of a damage tolerant design. The factors to be taken into account for life extension, are however the same.

In-service experience is carried through to ensure any lessons learned are incorporated, the most recent being Issue 2 of Def Stan 00-970 (2). This not only incorporates changes to the fatigue design requirements for new aircraft but contains also specific guidance on what steps should be taken in preparation for aircraft life extension. Provisions for good fatigue design include materials selection and fatigue testing. The latter has the requirement to do two fatigue tests, one on a pre-production aircraft and a later test on the production standard. Once the fatigue clearance is complete, the evidence is documented in the Fatigue Type Record.

A notable feature is an increase in loading severity in the longer term in service. A factor of 1.2 is applicable for example, in wing bending, due to increases in aircraft mass and severity of the normal acceleration spectrum. It is mandatory to carry out a tear-down inspection, both of the fatigue test article and of in-service aircraft on an opportunity basis.

Comparison of in-service usage with design assumptions is fundamental to airworthiness in determining equivalent safe lives and inspection criteria. With particular reference to modifications and concessions, there is a need for improved traceability of records.

Continued structural integrity for the RAF VC 10 was the subject of a paper from British Aerospace Military Aircraft and Aerostructures which concerns an early example of the fail-safe philosophy. After 1988 a series of events combined to prompt a review of VC 10 structural integrity, the original limits would be disregarded, and re-assessment applied against the criteria of a wholly damage tolerant regime.

It was apparent that at this time, the fleet was flying ahead of the safe lives derived for the wing and fuselage, which was borne out by in-service defects. Corrosion would cause further problems and the structural integrity had to be underwritten by inspecting all areas.

A range of short- and long-term measures was implemented. In the former, inspection intervals were assessed to account for the higher RAF operating weights and the Structural Sampling Programme was refined to include a much greater level of directed inspection. The long-term solution was to change the structural integrity philosophy of the aircraft to wholly damage tolerant and to produce a complete re-analysis of the airframe using a finite element model (FEM).

A tail plane refurbishment programme was introduced to counter the growing evidence of corrosion in this area and a programme of proof pressure testing was undertaken. It was necessary to create five complete aircraft VC 10 FEMs representing the five variants in service with the RAF (Figure 4). Validation of the FEM against two main sources of data was achieved. These sources were existing stress data detailed in BAC Weybridge reports originating from the VC 10 certification test programme and full airframe static tests carried out specifically to validate the Fatigue Type Record (FTR) FEM. The FTR identified many primary damage calculation sites across all the variants. A quantity of data from the operational load measurement (OLM) programme was available.

Figure 3 Development of damages during test progress - centre fuselage/wing specimens

Several exercises were subsequently undertaken; for wing load measurement; fin load measurement; and undercarriage load measurement. Various changes were incorporated and in 1998 a reappraisal was initiated which is expected to lead to a new formal Structural Integrity philosophy for the VC 10 defined by a branch of the RAF at Wyton, assisted by the DA and DERA. The post-FTR work will not be completed before 2000 and emphasises the difficulties encountered in altering the original fail safe design principles to a safer, albeit more tortuous, domain of damage tolerance.

Also from British Aerospace Airbus came a paper on the interacting effects of repairs in close proximity on the aircraft structure. The project sets out to establish the design philosophy for use in commercial aircraft for a variety of common repair arrangements. As aircraft age and repairs are required at intervals, often in close proximity, stiffness, stress and reduction in life prediction will need to be reassessed. Other activities will also lead to proximity problems, such as installation difficulties.

Finite element analysis was carried out on various repair proximity arrangements and validated by using a three-dimensional photo-elastic model of the structure. The areas of wing structure examined were representative of the top and bottom booms of the front and rear spars, centre spar top and bottom booms to butt straps and skin. The types of repairs considered were: spar angles in close proximity to other spar angles and bushed holes; skin plates in close proximity to other skin plates; and a variety of bushed holes along a staggered pitch and in relation to spar stiffening members.

Figure 4 Correlation of VC 10 C Mk 1 FEM fuselage stresses with test results

The study results for the wing structure gave dimensions at which no proximity effect needs to be considered. Also, for the centre spar and front and rear spars, factors have been determined. For the centre spar, a table gives the factor to multiply the end bolt hole fatigue life by, to give the life which includes the proximity effect for both internal angle and external skin plate reinforcings. For the front and rear spar, a factor is given to multiply the spar boom fatigue life by, to give the life which includes the proximity effect for the embodiment of an internal angle attaching to the horizontal boom and vertical web. Both these tables give results depending on the distance between repair end bolts. Factors are also derived for bushed repairs.

For the fuselage structure, boundary element analysis was used to model a uniaxially loaded skin lap joint. Nine boundary element models were constructed with varying separations between a pair of cutouts; 5in, 6in, 10in, and 14in separation models were all tested to investigate proximity effects under hoop stress loading. Results indicated that 14in was the limit beyond which no proximity effect need be considered for the fuselage joint investigated.

This investigation has created an awareness of proximity effects on wing and fuselage structure for the application of repairs at various stages of an aircraft life cycle, which form multiple repair arrangements. Multiple repairs in close proximity increase fastener loads and local stresses particularly at reinforcing end bolts. Even small increases in stress levels can result in significant reduction in fatigue life and the repairs examined enable answers to be given for proximity dimensions.

Life extension

Details of the Tornado Life Extension Programme were provided by British Aerospace, which is based on a number of aspects. The programme attempts to identify any life-limited aspects of the current design and focuses on the probability of a safety relevant event occurring. The airframe having been designed to meet safe fatigue life requirements, the fatigue life of the structure is proved by testing. This is achieved by a programme that subjects a representative airframe to the duty cycles defined within the requirement document.

Changes over the past 20 years in the world have changed the environment and duty cycle to which the aircraft operates. The fatigue index compiled from flight data, denotes the proportion of the safe fatigue life consumed. Fatigue sensitive areas are monitored structure. In the 1980s the Operational Load Monitoring (OLM) was established to check the accuracy of the fatigue meter formulae, provide for adjustment, and to enable fatigue life to be determined for parts of the airframe not able to be monitored by means of the fatigue meter. Emphasis is also placed on the control of corrosion.

For extending the life of systems and equipment, the strategy follows that of clearance for the structure although the process is quite different. Four phases are in being. First, a hazard analysis is undertaken and safety relevant equipment identified. Then it is determined whether to replace an item or pursue a re-qualification programme. The third phase involves suppliers undertaking the re-qualification activities, and the fourth phase addresses the aircraft level of certification and an update to the life items list, which forms part of the aircraft's documentation suite.

On the engine the Design Authority (Turbo Union) undertakes a detailed Flight Safety Review. This combined a theoretical assessment of potential failure modes of all aspects of the engine coupled with a comparison with in-service experience. Both the potential impact and the probability were assessed and a hazard matrix prepared. So called Group A parts, essentially the high energy rotating parts, have their lives monitored and an established method is available for changing those lives.

The embodiment strategy for life extension seeks to minimize life cycle costs and maintain aircraft availability. Where this is not possible, however, the Services will need to re-provision. Their ability to do this is a function of spares availability and obsolescence issues.

A lesson learned is the need to prepare for life extension well in advance of when it is needed. It is also certain that one of the impacts of a life extension programme will be a reduction in aircraft availability. It is recognised that the Services are constrained by aircraft with specific capabilities. The time involved in refurbishment, etc., is also a major consideration. In addition, accurate tracking of life consumed for life extension programmes needs to be expanded, by log carding. Also, any life extension programme should be treated as a project in its own right, from the outset.

Structural life extension of the RAF Hawk was described, also by British Aerospace, with the main components of the programme being four tasks. First, is the installation of a new wing standard to the Full Scale Fatigue Test (FSFT) and its continuation to the Life Extension Fatigue Test (LEFT). Second, the assessment of items and components not fitted to the FSFT, such as the aileron which was tested separately and the flaps which had been cleared analytically. Third, the assessment of build concessions (around 5,000), in-service modifications and repairs which were cleared against the original design requirements and, again, were not fitted to the FSFT. Fourth and last, the development of modification, inspection and replacement strategies is required.

The Hawk is operated in three different roles in the RAF, with variation in operational usage together with clearance factors, having significant effects on the life extension programme. The Hawk is currently the subject of an Operational Load Measurement Programme to establish the in-service loading of major structural components. It began in 1996 and is due to end in 2000. Certain components were identified as requiring further action as the life extension capabilities of these was not cleared by the LEFT. This was approached in a number of phases.

Two examples of non-LEFT component clearances were given; the pressure cabin, and the airbrake. The pressure cabin assessment considered 31 features in way of 19 locations, all of which met the life extension requirements (factor 10). The original fatigue clearance of the airbrake was performed in a similar manner to the pressure cabin, based on a static analysis employing fatigue "superfactors". This was not considered sufficient to support the life extension of this structure and a separate fatigue assessment was undertaken.

During the development of the life extension targets and the performance of the non-LEFT structure assessments, the LEFT has continued to extend the clearances of tested structure and consequently identified other problems.

In addition, and following the discovery of MSD being identified in the fuselage fuel tank side skins, what was originally intended to be modifications swung in favour of replacement of the rear fuselage and 80 of these are currently being supplied.

From the University of Auckland came a paper on Life Extension of Impact Damaged Honeycomb Sandwich Panels. This is becoming increasingly important and the study analyses minimum gauge, non-metallic honeycomb wing panels subject to impact damage. The damage is caused by "soft body" impactors travelling at elevated velocities to simulate bird strike and other soft debris.

Based on the findings of this study, some guidelines for the repair of soft body damage have been developed. A proposed repair technique should have the following features. It must prevent the formation of compression wrinkles across the impact defect; inhibit the growth of compression wrinkles; reduce the cyclic manipulation of crushed core material which further lowers the residual properties of crushed core; take into consideration the dramatic load capacity reduction that may be encountered with BVID (barely visible impact damage); the load reduction is relatively insensitive to defect size; and, allow effective shear transfer to the compression face sheet in the defect area so that the compressive load may be transmitted.