11:00   Life Extension and Management of Ageing Fleets II
Chair: Thierry ANSART
30 mins
Slawomir Klimaszewski, Andrzej Leski, Krzysztof Dragan, Marcin Kurdelski, Miroslaw Wrona
Abstract: In this paper selected problems connected with developing Helicopter Structural Integrity Program (HSIP) of Polish Army Mi-24 Hind Helicopters have been presented. The problems include: 1. Collecting, analysing and forecasting of operational helicopter’s failures 2. Developing of FE model of helicopter structure 3. Operational loads monitoring 4. Fatigue life estimation 5. Implementation of modern NDE methods In the paper abilities of IT system, the SAMANTA system used in Polish Armed Forces for collecting and analysing operational and maintenance related information have been discussed. Advantages from application of relational data base technology implemented in the SAMANTA system and based on the Oracle platform have been shown. In order to develop FEM model of Mi-24 structure reverse engineering (RE) methods have been used. The RE methods made possible helicopter structure to be digitally reconstructed and transferred to FEM software. Two of RE methods have been used: a) a digital photogrammetry method by means of TRITOP system, b) 3D scanning by means of ATOS III system. Service loads spectrum is obtained from operational loads monitoring program. The stresses and loads at selected points where measured during several test flights. Moreover the procedures of helicopter structure fatigue life estimation based on safe S-N curve concept and crack growth concept has been included in the paper. A modified NDE/NDI program is developed in order to support HSIP. Special attention was paid on detection of subsurface defects such as hidden corrosion and cracks by means of DSight Aircraft Inspection System (DAIS II), Magneto-Optic Eddy Current Imager (MOI-308) and MAUS V C-scan system. Above mentioned HSIP was developed by Polish Air Force Institute of Technology under research project founded by Polish Ministry of Science and Higher Education and Polish Ministry of National Defence.
30 mins
Manfred Heller, Madeleine Burchill, Ron Wescott, Witold Waldman, Robert Kaye, Rebecca Evans, Marcus McDonald
Abstract: Typically during the life of an airframe, a few key stress-concentrating locations can become fatigue critical, and an effective repair option is needed. Such repairs can provide significant economic benefits, by avoiding the need for component replacement, as well as usually increasing the interval between costly periodic in-service inspections. The Defence Science and Technology Organisation Australia (DSTO) has developed a unique life extension approach, where optimised rework shapes are designed and used. Two distinct scenarios for optimal shape reworking are: (i) a repair, where damaged material is removed and (ii) as a pre-emptive repair measure to avoid cracking. In both cases, apart from removing material, the optimised shapes provide significantly reduced stress peaks, thus leading to increased fatigue life, inspection intervals and airframe availability. The optimisation approach is based on an analogy with biological growth, and is implemented as an iterative gradientless finite element procedure. The optimal shapes are unique and depend on local loading and practical geometric constraints. An achievement of the DSTO work has been the transition from theory and generic developments to practical applications. Here, lessons learnt from the practical experience have also been used to enhance the theoretical methods, leading to an improved practical capability. Hence in the present paper, we cover the key issues relating to the theoretical and practical developments for rework shape optimisation, as follows: (i) Implementation of fully automated numerical methods, in 2D or quasi-3D. Advanced features include; minimising the magnitude of the multiple, constant-stress segments around the stress concentrator boundary (i.e. holes); robustness to account for perturbations in the direction of the dominant loading, or multiple load cases; and geometric constraints such as minimum radius for manufacture. (ii) Key results from generic studies and benchmark analysis, covering both unique and transferable shapes dependent on loading and geometric conditions. (iii) Practical applications and demonstrators including lessons learnt. This focuses on open holes and section runouts, including applications for F-111 fleet aircraft, F-111 fatigue test articles, and F/A-18 and P-3C components. (iv) Fatigue life effects due to the reductions in peak stresses and stress intensity factors, relating to the relevant airworthiness philosophy, i.e. safe life or safety by inspection. Taking into account the effect on fatigue life of the following; minimised peak stress, manufacturing constraints, robustness constraints, and non-destructive inspection limits. (v) Post processing the initial numerical shapes designs for effective computer aided manufacture. (vi) Improved manufacture of optimised shapes, particularly focussing on simplified in situ manufacture of shapes using compact jigs and semi-automated tooling. (vii) Approach to transitioning and certification of such repairs, and how optimised shapes themselves can be simply repaired should re-cracking occur (although unlikely). This includes recommendations on the effective use of the technology based on experience gained. (viii) Relevance of the technology to initial design, with an example from a combat aircraft. In summary, this paper presents the key achievements from an extensive work program undertaken by DSTO under the sponsorship of the Royal Australian Air force over the last twelve years. Highlights include: enhancements to the optimisation method, development of novel in-situ machining techniques and the successful implementation for RAAF aircraft. As indicated above, the paper is very much in tune with the conference philosophy of ‘bridging the gap between theory and operational practice’.
30 mins
Robert Rutledge, David Backman, Roger Hiscocks
Abstract: In-service component damage on the wing fold shear tie of a Canadian Forces aircraft, led to a full scale fatigue test using loads provided by the repair and overhaul contractor. The objective was to collect crack growth data at the wing-fold aft-spar shear-tie and to determine appropriate inspection intervals. Prior to fatigue testing, a strain survey was carried out and an automatic continuous ultrasonic inspection of the test area was developed and linked to the control system to stop the test after the development of a crack ranging from 0.015 to 0.020 inches. After several controlled stoppages, indicating that this size of damage was present in the critical location, the test was stopped and the paint covering the area was stripped to inspect it with liquid penetrant to confirm the ultrasonic indication. This inspection confirmed that an appropriately sized crack, typical of damages found in-service, was present. To remove this damage in the in-service aircraft the repair and overhaul contractor planned to blend this part. Prior to blending the test specimen the crack was carefully excised for examination. The excised pieces were examined using both optical and electron microscopy and quantitative fractography was used to develop a crack growth curve estimate of the in-service pre-modification crack. After excision of the damage, the proposed fleet structural modification planned for all wings was introduced by repair technicians. This hand blended modification of the geometry changed the component strain distribution and high first principal strains were found under peak loads at test restart. A digital image correlation system that used white light photogrammetry principles was used to measure both surface strain as well as the altered geometry in the blend region. This geometry data was provided to both the client and the repair and overhaul contractor for finite element analysis of the geometry modification. Based on the repair and overhaul evaluation, the applied spectrum peak load was reduced by approximately 8% to be consistent with the intended blend geometry peak strain/stress. The post modification phase of the fatigue testing was then started. Cracks developed in the new critical area and these coalesced into a single dominant crack. The cracking was monitored automatically using the digital image correlation system. The same system was also used to measure peak strains periodically during the fatigue testing. Testing applied the equivalent of seven life times of simulated flight hour usage in the two phases of the fatigue testing. This included the required factor of five lifetimes, post-modification, for an unmonitored dynamically loaded component. The applied spectrum alternated between blocks of loads for a tip mounted AIM-9 with fins-off and fins-on. By the end of the test a significant size crack had developed. To complete the test, a residual strength test was carried out to failure. Post test quantitative fractography was used to obtain the crack growth rate information. This data and the residual strength test failure load were then used to generate a residual strength curve. Based on the loading spectrum, a Gumble extrapolation was carried out to establish the one in six thousand hour design limit load. This was then scaled to obtain the residual strength, end of life, 120% load and an end of life critical crack size was estimated. The client and repair and overhaul contractor utilized this crack data to determine the appropriate inspection intervals.