Root cause


The first generation of composite primary structural components are now entering the twilight of their certified service lives. Life-extension efforts are underway for many affected aircraft, and efforts are primarily geared to the continued safe operation of the metallic components, for which a fatigue life can be quantified. However, many of the composite components do not have a known measurable fatigue life, so any extension of their lives cannot be accomplished via the same methodology.

Scientists at the National Institute for Aviation Research (NIAR) at Wichita State University are investigating aging of the McDonnell Douglas (now Boeing) F/A-18 Hornet’s wing structure, which consists of AS4/3501-6 carbon/epoxy composite wing skins and composite-to-titanium step-lap joints bonded with FM-300 film adhesive at the wing root and pylon fittings. The research program is funded by the US Office of Naval Research and monitored by the Airframe Technology Branch of Naval Air System Command (NAVAIR) – Air Vehicle Department.

The research team has formed collaborations with international F/A-18 users, such as the Royal Canadian Forces and Defense Science Technology Organization (DSTO) of the Royal Australian Air Force, to expand the use of data gathered during this effort. This program provides a quantifiable, risk-based assessment methodology for determining the capability for life extension in composite structures, combining original certification and operational use methods.

The research program is designed to address one of the biggest concerns with an aging aircraft fleet – the unknowns that emerge with little or no warning, raising the concern that an unexpected phenomenon may suddenly jeopardize an entire fleet’s flight safety, mission readiness and/or support costs.

F/A-18 wing root

The F/A-18 wing root ‘step-lap joint’ is one of the best examples of a bonded primary structure certified and deployed on an aircraft in the USA. Since a wing root bonded joint transitions from the composite wing skin to a titanium fitting for attachment to the fuselage, it is a complex joint.

The research program first focused on evaluating the residual strength of the composite-to-metal wing-root joint area after a lifetime of aircraft service and assessing the service life remaining based on the history of use. More than 60 25in-long tapered dog bone test specimens were extracted from eight decommissioned F/A-18 wing skins. A reliability analysis was conducted on residual strength and fatigue life data to compare against the original certification data.

Spectrum loading, representing fleet use, is used for cyclic testing to determine the remaining life of the step-lap joints. The residual strength tests conducted on elements extracted from decommissioned wing skins after one service life indicated that the service history, including environmental exposure, had no significant effect on the residual strength of the bonded joint.

Furthermore, the additional spectrum fatigue tests in a laboratory environment indicated that the remaining life of the joints was substantial and the residual strength was unaffected by the additional fatigue lifetimes. Prior to fatigue testing, some of the elements were exposed to an extreme salt-fog environment to simulate the exposure to aircraft carrier-based operations. Based on the testing conducted during this phase, the elements had more than five additional fatigue lifetimes remaining under a tension-dominant fatigue spectrum, and several specimens survived 10 lifetimes without significant residual strength degradation.

During tension-dominant spectrum fatigue loading, which is more critical on metallic components, a corner crack in titanium was observed in one of the step-transition areas. This crack propagated first through the thickness of the titanium and then across the width. Although the titanium was completely severed and the crack continued into the composite and propagated as a large delamination, the remainder of the joint continued

to carry loads for a large number of fatigue cycles before overloading the adjacent structure and resulting in final failure.

When subjected to a compression-dominant fatigue spectrum, which is more critical on composite parts, specimens survived 30 lifetimes of additional fatigue cycles with no significant residual strength degradation to the parts.

Non-destructive inspection

During fatigue tests, the specimens were non-destructively inspected using ultrasonic scans and pulse thermography equipment. In addition, traveling microscopes were used to monitor microcrack formation in the joint area and a photogrammetry image-correlation technique was used to monitor full-field displacements and strains.

Specimen compliance was monitored periodically to detect potential stiffness loses due to damage growth. Non-destructive inspections around the joint area were instrumental in understanding the microcrack formation of composite details near the joint area, enabling inspectors to interpret similar findings in the field. Fatigue data indicated that despite the formation of microcracks, the joints were able to carry fatigue loads and showed no change in stiffness. A detailed failure analysis and an investigation into findings from non-destructive inspections were carried out to understand the progressive failure mechanism.

F/A-18 inner wing

Upon completion of the fatigue life assessment of the step-lap joint, the research was expanded to evaluate the remaining life of the F/A-18 inner-wing, which consists of composite skins.

Test articles consisted of center fuselage, inner-wings and trailing-edge flaps. The center barrel section of the fuselage is used as part of the inner-wing test fixture to ensure proper load transfer at the wing-root lugs. Simulated inboard leading-edge flap and the outboard wing are attached to each inner-wing box for fatigue-load application. This structure provides an opportunity to look at service life management and offers a useful insight into the sustainment of the active fleet.

Current methods of certification for aircraft composite structures rely on the development of safe use through fatigue testing. Since composite structures are designed to withstand environmentally compensated static loads with considerable analytic reductions in strength, it is rare that the full-scale fatigue testing of aircraft components demonstrates the capabilities of the composite structural members. In addition, the expense of fatigue testing rarely permits continued testing past the original design goals of the program.

These factors combine to prevent composite structures from failing during the fatigue test. As a result, there is little capability over the course of the aircraft’s life to relate in-service events to known fatigue limitations of the original certification test, and no mechanism by which to employ engineering principles for the extension of life.

The ability to use end-of-life aircraft structural components has been shown to be beneficial in many instances for the support of the existing fleet and provides a proactive approach to fleet maintenance.

Next phase wing tests

During the first phase, the wing structure will be cycled for an additional lifetime of test loads while monitoring for potential fatigue-induced damage through various types of health-monitoring techniques. This full-scale test program also provides an opportunity to evaluate current field-inspection techniques and structural health-monitoring systems for detecting known or suspected damage threats found during teardown inspections for the extended life period, as well as assessing novel inspection techniques to detect damage threats in a controlled environment.

The second phase will focus on damage tolerance of the structure to avoid catastrophic failure due to fatigue, corrosion or accidental damage throughout the operation of the aircraft. Low-energy, barely visible impact damage (BVID) to large battle damage will be simulated and monitored during continued fatigue cycling to assess damage containment and/or the rate of growth so that the assumptions used during original certification related to damage tolerance philosophy and failure mechanisms are further validated. Subsequently these damages will be repaired and the durability of the repairs will be evaluated through further fatigue testing. An accidental crack detected on the aluminum leading edge of the right trailing-edge flap was repaired following a procedure documented by US Air Force Research Laboratory.

This composite ‘wet lay-up repair’ was subjected to pulse thermography and photogrammetry (image correlation) full-field strain evaluation prior to spectrum fatigue loading, and it was found that the localized displacements and the far-field strain at the periphery of the composite repair patch remained consistent with measurements prior to the repair. In addition, periodic inspections during one lifetime of fatigue testing of the trailing-edge flap indicated that the composite wet lay-up repair successfully prevented further damage growth without adversely affecting the local stiffness of the structure.

F/A-18 conclusions

The results of this test program will help to determine the airworthiness of the F/A-18 airframe and to assess any need for preventive maintenance for mitigating risks beyond its original design life. Lessons learned from this research will provide insight into the aging aspects of other non-metallic aircraft structures and influence the use of advanced materials on new aircraft being proposed for military service as well as maintenance of the existing fleet.

The full-scale structural test provides the opportunity to look at service-life management and gives useful insight into the sustainment of the active fleet. This research program will provide useful information on service life assessment, which assesses assumptions used in establishing the remaining service life, as well as on service life extension, which develops structural modifications to increase service life.

Teardown inspection

Upon completion of the full-scale fatigue testing, teardown inspections of critical areas will be conducted to assess the condition of the structure and compare the findings of previous teardown inspections conducted by the US Navy to identify the extent of known or suspected damage threats as well as any unexpected damage threats resulting from the service life extension. Furthermore, the post-test teardown findings will provide vital data to validate the NDI findings during durability testing and identify equipment capabilities and limitations.

Dr Waruna Seneviratne currently serves as a technical director at the National Institute for Aviation Research (NIAR) at Wichita State University and is the lead scientist in the aircraft airworthiness and sustainment research program

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