DAMAGE TOLERANCE CAPABILITY

This paper is dedicated to my friend and colleague Professor Dr Jaap S chijve, required to retire like many good aircraft components accordin g to a 'safe life' criterion but who is still young in mind, body and spirit, has not yet reached his unfactored endurance limit, and who co uld continue subject to a needed change in philosophy to 'retirement f or cause'. Professor Schijve has dedicated his life to aviation safety through his outstanding continuous research into aircraft fatigue and fracture phenomena. The Federal Aviation Administration, whose primar y goal is aviation safety, is indebted to Professor Schijve for his co ntinous counsel through many contributions to the literature in his su bject. The purpose of this paper is to encourage manufacturers of futu re transport aircraft to retain the large damage-tolerance capability designed into the first wide-bodied aircraft and to modify their metho dology to establish inspection thresholds for those structures incapab le of sustaining large obviously detectable damage. In the current eco nomic environment there appears to be a general trend to lower the lev el of safety built into the original wide-bodied aircraft to reduce as sembly costs and weight. A number of examples are provided highlightin g the implications of this general trend. Most of the wide bodied airc raft are designed to sustain a two-bay skin crack with a broken centra l stiffener at limit load. Some manufacturers would prefer to relax th is criterion for wide expanses of basic structure with a view to savin g weight. The implications of this are that extremely sophisticated ND I will be required over wide areas to satisfy the damage tolerance req uirements, thus creating a considerable burden on the operator. This i s illustrated by example. Most of the large transport aircraft manufac turers establish the threshold for detailed inspection of principal st ructural elements through a fatigue evaluations process without consid erations of initial manufacturing flaws. These thresholds are often as long as three quarters of the aircraft design life goal. This practic e has been thought to be satisfactory for fail-safe crack arrest struc ture. In this case a second line of defence exists in the event that a n initial manufacturing flaw nucleates into a propagating fatigue crac k during the service life of the aircraft. Since crack arrest structur e is usually capable of sustaining large obvious damage it is likely t hat such large damage would be detectable. However, principal structur al elements do exist that are incapable of sustaining large obvious da mage where initial manufacturing flaws could grow critical prior to th e threshold established by fatigue evaluation. This is illustrated by a typical example. Structures operating beyond the life substantiated by full-scale testing may be prone to multiple-site damage (MSD). Extr emely small in-service undetectable MSD has the potential for substant ially reducing lead crack and discrete source residual strength. This is illustrated by typical examples. The majority of large transport ai rcraft developed in the USA have circumferential crack-stopper straps attached directly to the fuselage skins to guard against explosive dec ompression failure in the event of undetected fatigue damage or discre te source damage. This was thought necessary after the Comet disasters in 1954. There is a current trend to eliminate these crack stoppers f or future designs and depend only on shear clips for crack arrest capa bility to save on assembly costs. The wisdom of this trend is challeng ed by considering examples and citing a number of secondary effects th at may prevent arrest of fast fracture.