Structural Integrity of Aging Airplanes

Non-destructive testing techniques, as they are practiced in the field of quantitative non-destructive evaluation, are at the basis of a comprehensive approach to secure the safety of aging aircraft. The applications, advantages and disadvantages of the principal NDE techniques are summarized in this paper. It is discussed that measurement models for these techniques, in conjunction with the POD concept, scanning plans and methods of graphical display, facilitate the selection of optimal procedures for specific inspection problems. These models also suggest NDE standards and calibration techniques and they can be an important part of inspection system validation and operator training. Four major components of a comprehensive QNDE program for aging aircraft are identified and briefly discussed.

This paper reviews some simplified computational strategies for the integrity analyses of fuselage panels in aging airplanes. The topics covered include: (i) a “direct stiffness” method for stiffened panels (with fastener flexibility being accounted for) with multiple cracks, using the alternating technique, (ii) MSD near a row of fastener holes, (iii) analysis of cracks with bonded repair patches, (iv) weight functions for multiple cracks, and (v) bulging near crack-tips in pressurized fuselages, and equivalent domain integral methods for computing fracture parameters for bulged cracks.

To date, the Air Force has applied the durability and damage tolerance requirements of MIL-STD-1530A in three areas: a) designing new aircraft; b) evaluating the durability and damage tolerance characteristics of aircraft which were designed prior to the current requirements of MIL-STD-1530A; and, c) evaluating structural repairs and modifications. These applications have used fracture mechanics principles in a deterministic manner. That is, flaw growth was predicted using a fixed potential flaw size, a fixed da/dn vs AK relationship, and a stress spectrum derived from a predicted average usage. While it was realized that there are many stochastic elements in the initiation and growth of cracks, the applied process was considered conservative. Initial flaw size assumptions were generally severe, tracking programs accounted for variations in usage severity, and inspections (if necessary) were scheduled at half the time required for the specified initial flaws to grow to a critical size. At the time of the assessments, there was a high (but unquantified) degree of assurance that fatigue failures and widespread cracking would not occur within the design operational lifetime.

This paper will discuss from a regulatory viewpoint some of the key technical matters in the ageing aircraft issue. Attention is focused primarily on research topics of interest, for an overview of the CAA approach to ageing aircraft in general the reader is referred to Williams [1] and James [2]. The topics covered here are

This paper presents a brief review of the Damage Tolerance Requirements (DTR) for civil aircraft, and of their impact and significance for both newly developed and ageing aircraft. Some improvements to the DTR are suggested. In principle, fracture control of aircraft is ensured by non-destructive inspection, although destructive inspection has been suggested for certain ageing aircraft, as will be discussed briefly as well.

The effect of prior exfoliation corrosion on the fatigue behaviour of 7178-T6 and 2024-T351 aluminium alloys is investigated. The fatigue crack growth rate in both as received and prior corroded material is measured in both wet and dry environments. The exfoliation corrosion is shown to enhance the fatigue crack growth rate in the 7178-T6 material whereas the same effect was not observed in 2024-T351.

Damage Tolerance Philosophies are being used today in the Design, Development, and Life Management of USAF turbine engines. The establishment of a Damage Tolerance Philosophy and implementation of a Fracture Control Program has been an integral part of the USAF turbine engine development process since 1978. It has been applied to new engine programs as part of the USAF’s Engine Structural Integrity Program (ENSIP) described in military standard 1783 issued formally in 1984. This military standard was reviewed and approved by the Aerospace Industries Association of America, Inc (AIA) in 1982. Damage Tolerance philosophy has also been applied to existing inventory USAF engines through structural durability and damage tolerance assessments. In all, fracture control programs have been applied to the F-100, TF-34, F100-PW-220, F100-PW-229, F110-GE-100, F110-GE-129, F101-GE-102, F109-GA-100, F-119, F120 and T406 engines and have resulted in the generation of NDE flaw size quantification probability of detection data and in implementation of enhanced nondestructive evaluation methods (i.e., eddy current) at manufacturing and at field/depot. Experience has shown that these inspection have been successful in detecting early cracking and in accelerating corrective actions. It has been highlighted that several development efforts in the last five years have been identified FPI process improvements that must be implemented within industry and Air Force depots to improve flaw detection reliability. The need to quantify detection reliability for imbedded defects is also identified. The engine development process has been evolutionary in terms of application of upgraded requirements. The new process of fracture control, sometimes referred to as damage tolerance, is contained in ENSIP and selects material as well as governs design and life management. Recent experience clearly demonstrates that the damage tolerance requirement is cost effective when assessed on a life cycle basis.

The concept of pressure proof testing of fuselage structures with fatigue cracks to insure structural integrity was evaluated from a fracture mechanics viewpoint. A generic analytical and experimental investigation was conducted on uniaxially loaded flat panels with crack configurations and stress levels typical of longitudinal lap splice joints in commercial transport aircraft fuselages. The results revealed that the remaining fatigue life after a proof cycle was longer than that without the proof cycle because of crack growth retardation due to increased crack closure. However, based on a crack length that is slightly less than the critical value at the maximum proof stress, the minimum assured life or proof test interval must be no more than 550 pressure cycles for a 1.33 proof factor and 1530 pressure cycles for a 1.5 proof factor to prevent in-flight failures.

Economic and market conditions have resulted in the use of commercial jet airplanes beyond their original economic design life objectives. The average age of the world airline jet transport fleet has increased from 8 to 12 years since 1980. Standard Boeing practices to ensure continuing airplane structural integrity include inspection and overhaul recommendations contained in maintenance manuals 2nd service bulletins. As airplanes exceed their economic design life objectives, the incidence of fatigue increases and corrosion may become more widespread. This presentation is focused on principles of durability and damage tolerance technology standards suitable for large teams of structural engineers; major tests and teardowns of airplanes retired from service; results from worldwide surveys of over 80 aging fleet airplanes and joint Boeing, airline and airworthiness authority reviews of service bulletins, corrosion control programs, basic maintenance and supplemental structural inspection programs, and structural repair quality. These initiatives have provided timely preventive maintenance recommendations that will permit continued safe operation of aging jet transports until their retirement from service for economic reasons.

NASA has initiated a research program with the long-term objective of supporting the aerospace industry in addressing issues related to the aging commercial transport fleet. The interdisciplinary program combines advanced fatigue crack growth prediction methodology with innovative nondestructive examination technology with the focus on multi-site damage (MSD) at riveted connections. A fracture mechanics evaluation of the concept of pressure proof testing the fuselage to screen for MSD has been completed. Also, a successful laboratory demonstration of the ability of the thermal flux method to detect disbonds at riveted lap splice joints has been conducted. All long-term program elements have been initiated and the plans for the methodology verification program are being coordinated with the airframe manufacturers.

This report concerns some of the factors relating to the Aloha Airlines accident which occurred on April 28, 1988, in the Hawaiian Islands. What is different about this occurrence is that in the past it has usually been the high fatality spectacular catastrophic accidents that have worked as the catalyst for change. High visibility aviation accidents with great loss of life would serve as the trigger for stampeding the government and industry into action. This phenomena is sometimes referred to as the “tombstone effect.” Examples of this premise are: (1) the Eastern Airlines flight 66, B-727 approach accident that occurred at the JFK International Airport on June 24, 1975, killing 113 passengers and crew. Although thunderstorm/down draft accidents, certainly, were not a new phenomena at that time, this accident because of the high visibility focused on it precipitated the first full scale efforts by the FAA and the industry to develop and implement windshear detection and training programs, and (2) the spectacular PSA mid-air collision accident over the city of San Diego on September 25, 1978, which claimed 144 lives. While, certainly, not the first mid-air collision accident, resulted in massive changes to the National Airspace and Air Traffic Control System including a new requirement for a network of terminal control areas which are still in place today.

In the past few years there has been much discussion about the design of fuselage longitudinal skin splices. This paper presents one manufacturer’s approach to the design of those splices. As such. this paper is not meant to be an authoritative source on the design of such details, but should be viewed as a case history of one way that has proven successful. This paper also details where research is needed to improve the design processes and the subsequent in-service inspection requirements for fuselage structure.

The types of multi-site damage which are observed in fuselage skins preclude use of common simplifications employed in simulating crack growth. Interaction of cracks with each other and with structural features causes arbitrary growth patterns. Representation of a number of arbitrarily growing cracks in a finite element model is a problem which can be addressed with new codes built on a topological data structure and employing automatic remeshing. Examples of such codes applied to problems of interacting cracks, crack interaction with rivet holes, and cracking beneath a boron/epoxy patch are presented. A technique for incorporating probability methods into simulations of arbitrary crack growth is also presented and discussed via an example problem.

Repairs using bonded composites have numerous advantages over mechanically fastened repairs. Adhesive bonding does not result in stress concentrations due to additional fastener holes. Composites are readily formed into complex shapes, permitting the repair of irregular components. In service damage monitoring is possible, with the appropriate fibre matrix system, by direct inspection through the repair using eddy current methods or by thermal emission measurements. This paper presents a bonded repair for fuselage lap joints containing multi-site damage. The effectiveness of this repair is confirmed by the results of a laboratory test program.

Multi-site damage (MSD) in the form of cracking at rivet holes in lap splice joints has been identified as a serious threat to the integrity of commercial aircraft nearing their design life targets. Consequently, to assure the safety of aircraft that have accumulated large numbers of flights, flight hours and years in service requires inspection procedures that are based on the possibility that MSD may be present For inspections of aircraft components to be properly focused on the defect sizes that are critical for structural integrity, fracture mechanics analyses are needed. The current methods are essentially those of linear elastic fracture mechanics (LEFM) which are strictly valid only for cracks that extend in a quasi-static manner under small-scale crack tip plasticity conditions. While LEFM is very likely to be appropriate for subcritical crack growth, quantifying the conditions for fracture instability and subsequent propagation may require advanced fracture mechanics techniques. The specific focus in this paper was to identify the conditions in which inelastic-dynamic effects occur in (1) the linking up of local damage in a lap splice joint to form a major crack, and (2) large-scale fuselage failure by a rapidly occurring fluid/structure interaction process.

The rapid crack propagation, crack curving and arrest mechanisms associated with a pressurized, thin-walled ductile steel tubes are used to develop a model of axial rupture of an aircraft fuselage. This model is used to replicate axial crack propagation along a line of multi-site damage (MSD) and crack curving and arrest near a tear strap of an idealized fuselage.

This meeting is being held because of an accident which occurred nearly two years ago. The topics to be covered addresses issues highlighted by that accident.

Flat coupons were tested in the laboratory to determine a fracture criterion for link-up of fuselage lap splice multiple site damage at adjacent rivet holes. Experiments were performed on 0.040 inch (1 mm) thick 2024-T3 clad aluminum sheet. Continuous and riveted lap splice coupons were tested with simulated uniform (equal crack lengths) and nonuniform MSD, and the effects of notch sharpness were also studied. A net section stress criterion was found to provide excellent predictions of fracture for uniform MSD and uniform stress distributions. This same criterion provides conservative predictions for nonuniform MSD in uniform stress fields. An overload/cyclic stress experiment was also conducted to explore the pressurized proof test scenario of ensuring structural integrity.

One way to avoid multiple site damage in airframes of the future is to screen design details by means of coupon tests. A simple conceptual model for estimating the risk of multiple site damage is presented. The risk is expressed as the fraction of similar details which can be expected to form fatigue cracks at times close enough to each other to allow a fracture cascade after the damage has propagated. Examples based on an approximate representation of typical transport category fuselage skin properties and stress environments show that the risk becomes significant for details which exhibit much less fatigue life scatter than is observed in plain fatigue tests and for which the characteristic life is of the order of the service life. Results for confidence limits are also presented to show that at least 40 replicate details must be tested to obtain adequate estimates of scatter and characteristic life.

A series of fatigue tests were performed on riveted panels of clad 2024-T3 without epoxy bonds. Fatigue crack initiation occurred at the apex of the rivet hole chamfers. Transgranular fatigue crack growth by ductile striation formation occurred through the sheet. The fracture features at low, medium and high growth rates were examined with the SEM. Microscopic crack propagation rates as measured by fatigue striation spacings correlate with macroscopic crack growth rates observed. The fatigue crack growth rate is fairly constant over a length of 6 mm (0.25 in.) from the edge of the rivet hole, due to the fact that the stress intensity range is approximately constant in this region. Transition to fast fracture and unstable crack propagation is readily identified due to marked yielding of the cladding material.

The philosophy as developed by Fokker with regard to the structural maintenance is presented. The maintenance schedules have always included a part with a mandatory character: the Structural Integrity Program (SIP). In this document the stress corrosion aspect is also covered.

As a result of the first FAA aging aircraft conference a Fokker/Piedmont-USAir Task Group was formed for the F28, resulting in a maintenance review activity. The first meeting of the Task Group was in March 1989.

The objectives of the Task Group are presented.

The review is now in the final stage. The last meeting is scheduled in May 1990.

The paper discusses the results of the review as far as these are available and closes with a number of conclusions.

A new test facility for evaluating the fatigue and fracture strength of stiffened and jointed aircraft fuselage panels has been built at Foster-Miller. The test fixture can accommodate 40-deg curved sector panels of 120 in. in length and 58 in. in width. Internal pressure loading up to 20 psi and corresponding longitudinal and hoop restraining loads can be applied. Cyclic pressurization loads can be used for fatigue studies. The stress distribution in the full-size round fuselage structure can be accurately simulated in the central region of the test panel. Internal pressurization fluid can be air or water. The test facility is initially intended for (i) identifying causative factors contributing to multiple site damage in curved structures, (ii) generating fatigue data for comparing with the data from flat coupons, (iii) evaluating the residual strength of old aircraft panels, and (iv) validating the finite element based theoretical model currently being developed by Foster-Miller.

The objectives of the work reported in this paper were to assess the probability of detection of Multi-Site Damage (MSD) in fuselage lap-splices in aging airplanes. The results indicate that a reduction in the mandated period between inspections should be considered. Perforce, the model had to rely on NDI inspection reliability measured on samples devoid of MSD. A recommendation is made for a program to acquire such data for MSD.

All AIRBUS A/C have been designed with special consideration to the fatigue behaviour and damage tolerance aspects from the beginning of the seventies.

The AIRBUS philosophy and the essential evaluation methodologies, used for type certification, which were used for AIRBUS A/C designed prior to or after introduction of the damage tolerance requirements as specified by FAR 25 Amendment 45 and JAR 25.571 are presented. The calculation methods and the extensive full scale fatigue tests, used for certification, are described, as well as special investigations to assess the effect of possible corrosion on fatigue and crack propagation.

In addition details and the status of activities related to ageing A/C, such as the establishment of the supplemental structural inspection program and FAA initiated ageing A/C tasks, are explained. This includes initial results from research programs to study the effect of periodical overloads on the fatigue and crack propagation behaviour of pressurized fuselages as well as activities to optimize crack propagation calculation methods for fuselages using stress-time-histories.

Problems of providing the safety of transport airplanes with respect to the fatigue strength have been considered by Soviet researchers since the mid 1950’s. The adopted approaches basically correspond to those used all over the world, however some aspects of both technical and managerial efforts have different features. The main one is that the airframe is considered as an aging one from the beginning of the aircraft fleet service. At the present time this is a complete system, which has shown a high efficiency. The methodological and managerial base of the system and some scientific and technical aspects are discussed in this report.

After briefly describing the principles of frozen stress photoelastic and moire interferometric analyses, and the corresponding algorithms for converting optical data from each method into stress intensity factors (SIF), the methods are applied to the determination of crack shapes, SIF determination, crack closure displacement fields, and pre-crack damage mechanisms in typical aircraft component configurations.

The results of displacement compatibility analysis, representing a variety of repair doubler and lap splice configurations, are presented with a view to illustrating how structural repairs can degrade the fatigue initiation life and damage tolerance capability of primary transport aircraft structure. Examples show that fatigue initiation life is directly related to the peak loads induced in the first fastener rows at the edges of repair doublers. Design of repairs to an equal or better static strength capability and the associated static strength analysis will not normally highlight these peak loads which can result in considerable degradation of structural fatigue life. Critical fastener loads, based on displacement compatibility analysis accounting for fastener flexibility, are parametrically presented for a variety of skin and doubler thicknesses. Suggestions are made on how repair designs can be modified to improve fatigue initiation life and subsequent fatigue crack detectability particularly in the event of multiple-site damage (MSD). The importance of riveting quality during repairs, often not up to initial manufacturing standards, is discussed with respect to fatigue initiation life. A simplified but conservative method to generate crack growth curves is discussed with a view to easing the analytical burden for the small modifiers. It is hoped that this information, together with conservative fatigue Sn data, will help the many small repair and modification stations gain an appreciation of the fatigue and damage tolerance quality of structural repairs. It is pointed out in the paper that the FAA regulations were amended in December, 1978 for new transport category aircraft and in May, 1981 for aging transport aircraft to include a damage tolerance philosophy. This means that any repair to a transport category airframe, which may effect threshold, frequency and type of inspection of principal structural elements, must be evaluated for it’s damage tolerance capability.

This is a review paper on the study of aging aircraft structures conducted in Japan.