Aircraft Structural Integrity Research Articles

Managing airframe fatigue from corrosion pits – A proposal

Fatigue management is generally well understood as reflected in Aircraft Structural Integrity Management Plans, which in some cases considers environmental degradation prevention, however limited provision beyond find and fix exists for corrosion repair. Thus the repair of corrosion can be a major through life cost driver and an aircraft availability degrader. This find and fix approach exists largely because tools are too immature to accurately assess the structural significance of corrosion when it is detected. This work aims to provide a crack growth basis for the justification of allowing detected pitting corrosion to remain in service for a limited period.

On the potential of supersonic particle deposition to repair simulated corrosion damage

Previous constant amplitude laboratory tests have shown that supersonic particle deposition (SPD) can be used, as an alternative to mechanically fastened doublers, to enhance the structural integrity of aircraft structures. This paper compliments these findings via a numerical study into the potential for SPD scarf repairs to repair simulated corrosion in thin aircraft structural components. The results of this study, when taken in concert with the results of prior applications to full scale fatigue tests and laboratory tests, suggest that a SPD scarf repair to simulated corrosion damage subjected to load spectra representative of helicopters, fighter, maritime, and transport aircraft has the potential to give lives that are compare favourably with either the original design objectives or the stated operational life of the aircraft.

Supersonic particle deposition as a means for enhancing the structural integrity of aircraft structures

The paper presents the results of experimental studies into the application of supersonic particle deposition (SPD) for repairing and enhancing the airworthiness and integrity of aging aircraft structures. Presented are the results of coupon tests on the application of SPD doublers to mechanically fastened joints and to simulated corrosion damage tested under constant amplitude fatigue loads. These successful tests are then supported by an application to an F/A-18 Hornet wing attachment centre barrel laboratory test article.

Assessment of damage containment features of a full-scale PRSEUS fuselage panel

The most recent generation of integrally stitched composite technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept, which shows promise in enhancing structural integrity of aircraft and aerospace structures. An assessment of damage-containment capabilities of the PRSEUS concept for a full-scale fuselage panel was recently conducted under a joint program conducted by the National Aeronautics and Space Administration, the Federal Aviation Administration, and The Boeing Company. The panel was subjected to axial tension, internal pressure, and combined axial tension and internal pressure load conditions up to fracture, with a two-bay through-the-thickness notch severing the central stiffener. Panel deformation was monitored via strain gages, linear variable differential transducers, and digital image correlation. Damage formation, propagation, and accumulation were monitored visually throughout loading. Post-failure nondestructive and teardown inspections were performed in the vicinity of the notch tips where stable damage progression occurred. These test results demonstrated that the PRSEUS design exhibits improved damage-containment capability.

Standardized Test Method for Corrosion Pit-to-Fatigue Crack Transition for AA7075-T651 Aluminum Alloy

Corrosion damage (pit) is a stress raiser that can have deleterious effects on the fatigue life of airframe structural components. A better understanding and method for modeling the corrosion pit to fatigue crack transition would advance the fidelity of aircraft structural integrity estimates and fleet management decision making. Here, the focus is on developing a standardized fatigue test method for investigating the transition of a corrosion pit to fatigue crack in aluminum alloy AA 7075-T651 specimens. The standardized test method requires the development and validation of two sub-protocols (1) a pitting protocol to create a corrosion pit less than 200 μm diameter at the intersection of the central hole bore and planar surface of sheet and (2) a spot welding protocol for attaching the direct current potential drop (dcPD) probes on either side of the corrosion pit for fatigue crack growth measurement. A dcPD fatigue test method coupled with a unique 10-4-6 marker load sequence is used to measure the fatigue crack growth. The crack shape evolution and crack growth life are predicted using AFGROW.

Managing Fatigue from Corrosion Pits in Aircraft Structures

Despite corrosion prevention or protection schemes/treatments and corrosion prevention and control plans, in-service corrosion does occur and has the potential to impact the structural integrity of aircraft. Whilst the fatigue management of the aircraft is generally well understood as reflected in typical Aircraft Structural Integrity Management Plans (ASIMP), which in some cases contain environmental degradation plans, limited provision beyond find and fix exists for corrosion repair. Thus the repair of corrosion can be a major through life cost driver as well as an aircraft availability degrader. This find and fix policy exists largely because tools are currently considered too immature to accurately assess the structural significance of corrosion when it is detected. In this paper a process is described which should allow an alternative to the current find (corrosion) and fix philosophy for pitting corrosion. The method is intended to maintain a probability of failure consistent with ASIMP structural certification requirements for fatigue cracks initiating from corrosion pits for a specific period. Unanticipated maintenance costs significantly more than planned maintenance. Thus delaying the repair of pitting corrosion until the next scheduled maintenance, should save considerable resources and improve aircraft availability. The development of analytical tools capable of accurately assessing the effect of corrosion on the durability of a structure would be considered a major advance for the ASIMP.

A Modified Compliance Method for Automatically Measuring Fatigue Crack Growth <i>In Situ</i> during Spectrum Fatigue Testing

The fatigue critical structures of military aircraft are generally subjected to variable amplitude flight spectrum loading. Maintaining aircraft structural integrity to ensure safe operation of the fleet is critically dependent on accurate analysis and reliable prediction of fatigue crack growth in those structures under service operating conditions. To achieve this goal, laboratory experimental methods that can accurately measure and monitor fatigue crack growth under variable amplitude loading are required. This can be challenging as no test standard exists to guide the process of fatigue crack growth measurement under variable amplitude loading conditions to ensure the accuracy of the test results. This challenge was addressed by developing a modified compliance method as described in this paper. The results presented employed a modified compliance method complemented with a travelling microscope technique and marker band loads. The modified compliance method developed is able to measure in-situ, fatigue crack growth of standard compact-tension specimens under a fighter flight spectrum loading. The marker band loads and microscope readings were used to assist the post-test validation using quantitative fractography. The results from this study have demonstrated that the modified compliance method can produce consistent and accurate fatigue crack growth data under variable amplitude loading conditions.

Composites and MIL-STD-1530C

Purpose – The purpose of this paper is to discuss changes to MIL-STD-1530C “Aircraft Structural Integrity Program” to account for the increased usage of composites in aircraft structures. Design/methodology/approach – The evolution of the Aircraft Structural Integrity Program is presented and the five tasks that comprise the program are assessed for compatibility with composite aircraft structures. Findings – This paper identifies a number of recommended changes to MIL-STD-1530C to ensure that the unique behaviour of composites is considered within the Aircraft Structural Integrity Program. Originality/value – This paper recommends changes to MIL-STD-1530C to account for composite aircraft structures, thus providing assurance compatibility of the Aircraft Structural Integrity Program with composite materials.

Development of design charts for stress intensity factors at tips of multi-site cracks in unstiffened curved panels

Multi-site damage need to be addressed and evaluated in order to assess the integrity of aging aircraft structures. One of the problems recognized in the recent times is the effect of interaction between two or more cracks in the close neighborhood in such structures. The present paper deals with such a problem and presents numerical estimates of stress intensity factors at a crack tip in an un-stiffened curved panel with a secondary crack in the vicinity of a primary crack. The results are presented in the form of design charts. These results should be useful in evaluation in the damage tolerance evaluation of aircraft structures with multi-site damage.

Application of supersonic particle deposition to enhance the structural integrity of aircraft structures

Aircraft metal components and structures are susceptible to environmental degradation throughout their original design life and in many cases their extended lives. This paper summarizes the results of an experimental program to evaluate the ability of Supersonic Particle Deposition (SPD), also known as cold spray, to extend the limit of validity (LOV) of aircraft structural components and to restore the structural integrity of corroded panels. In this study [LU1]the potential for the SPD to seal the mechanically fastened joints and for this seal to remain intact even in the presence of multi-site damage (MSD) has been evaluated. By sealing the joint the onset of corrosion damage in the joint can be significantly retarded, possibly even eliminated, thereby dramatically extending the LOV of mechanically fastened joints. The study also shows that SPD can dramatically increase the damage tolerance of badly corroded wing skins.

Can pitting corrosion change the location of fatigue failures in aircraft?

In this paper we hypothesize that pitting corrosion can alter the location of fatigue failures in aircraft. This is in addition to the reductions in fatigue life it causes. The investigation began by developing a predictive Monte Carlo model of the fatigue behavior of low-kt specimens of the aluminum alloy 7010-T7651. The model’s fatigue life predictions were an excellent match to experimental fatigue life data from previous DSTO research for the same specimen geometry and material. The model was then used to predict the proportion of failures due to pitting corrosion as a function of the location of the corrosion strike. These predictions were compared to the results of an experimental trial of specimens which had been corroded at the same offset locations. The predicted and experimental proportions of failures were in close agreement. The paper concludes with a discussion of an integrated model of corrosion pit nucleation and growth followed by fatigue crack growth to failure. The purpose of this integrated model is to give the Royal Australian Air Force an end-to-end model to predict the damaging effects of pitting corrosion on aircraft structural integrity.

Verification of an individual aircraft fatigue monitoring system

Fatigue monitoring of airframes has developed over the decades to the stage where it is now incumbent for the certification of fighter type aircraft to incorporate a fatigue monitoring system. These systems typically collect operational data for the calculation of the airframe’s safe-life or crack inspection intervals. Many of these systems are complex, incorporating such features as data integrity checking, strain gauge calibration algorithms and damage calculation algorithms to name a few. Whilst it may be possible to validate the robustness and accuracy of specific system components (e.g. the damage algorithm can be tested against fatigue coupon results), the verification of the performance of the in-service system as a whole presents a much bigger challenge.In this paper, the verification of the Royal Australian Air Force’s F/A-18A/B Hornet individual aircraft fatigue monitoring system is outlined. The availability of a significant number of ex-service centre fuselage sections with known usage has facilitated this effort. Using an enhanced teardown procedure, in-service fatigue crack growth has been identified at a significant number of locations. All the in-service cracking corresponded to the same locations found cracked in the fatigue certification full-scale test article that was used to calibrate the usage monitoring system, so that by comparing the measured in-service growth with the test-demonstrated growth the functionality of the monitoring system could be assessed. This assessment should reveal the effectiveness of the system in providing robust fatigue life expended indices to help ensure that structural integrity boundaries are not exceeded. For this comparison, the crack growth was measured using quantitative fractography.It is believed that this work is the first example of using the crack growth in retired structure of known usage to verify a fatigue tracking system that incorporated significant aircraft structural integrity elements including tracking philosophy, structural fatigue lifing methodology, full-scale fatigue test results, design standard interpretation and retirement considerations.

Interaction of hybrid pressurised cylindrical structures subjected to blast loading

This paper studies the dynamic behaviour of fibre metal laminate pressurised barrels subjected to internal blast loadings using the multi-material arbitrary Lagrangian–Eulerian (MMALE) formulation within the explicit finite element software RADIOSS. The barrel set-up was based on well controlled and minimal dynamic fracture tests, which aim to emulate conditions onboard a commercial aircraft. The effects of length and static prepressurisation were investigated to determine the interaction between barrel deformation and reflected shockwaves. In this paper, it is shown that the length of the barrel may significantly induce severe deformation modes as a result of shock reflection waves from the top surface of the barrel. Finally, the effect of pre-pressurisation was proven to alter the structural response to an internal explosion by providing additional internal energy to propagate the blast wave. As a result, it has been concluded that prepressurisation of an aircraft fuselage will increase the structural damage due to the detonation of an onboard explosive device and should be included in all future studies to more accurately model the dynamic crack growth and structural integrity of typical aircraft structures.

Investigations on TC4-DT titanium alloy using laser shock processing with high energy

Laser shock processing or laser peening has become an advanced and promising surface treatment technique and has been shown to be effective in increasing the resistance of metallic components to high-cycle fatigue (HCF), stress corrosion cracking (SCC), wear, etc., through imparting compressive residual stress on the surface of a number of metals and alloys. The higher laser energy/power density could induce the deeper effect layer, larger action area, and better treatment effect. The property of damage tolerance has become the design concept and selection criterion of breakage-safety of meeting the requirements for advanced aircraft and engine structural integrity. The TC4-DT is two phases titanium alloy of damage tolerance with high fracture toughness and resistance fatigue extension through com

Exfoliation Corrosion and Pitting Corrosion and Their Role in Fatigue Predictive Modeling: State-of-the-Art Review

Intergranular attack (IG) and exfoliation corrosion (EC) have a detrimental impact on the structural integrity of aircraft structures of all types. Understanding the mechanisms and methods for dealing with these processes and with corrosion in general has been and is critical to the safety of critical components of aircraft. Discussion of cases where IG attack and exfoliation caused issues in structural integrity in aircraft in operational fleets is presented herein along with a much more detailed presentation of the issues involved in dealing with corrosion of aircraft. Issues of corrosion and fatigue related to the structural integrity of aging aircraft are introduced herein. Mechanisms of pitting nucleation are discussed which include adsorption-induced, ion migration-penetration, and chemicomechanical film breakdown theories. In addition, pitting corrosion (PC) fatigue models are presented as well as a critical assessment of their application to aircraft structures and materials. Finally environmental effects on short crack behavior of materials are discussed, and a compilation of definitions related to corrosion and fatigue are presented.

A Structural Health Monitoring of a Pitch Catch Active Sensing of PZT Sensor on Normal, Damage and Repair Aircraft Spoiler

At present, aircraft structural integrity is a concerned due to heavy usage of composite materials and cost saving on the operational. Structural health monitoring system is one of condition based monitoring introduced to supplement the current aircraft maintenance non destructive inspection. One approach is to embed or attach sensors such as lead zirconate titanate (PZT) to detect the anomalies either passively or actively. Due to the aircraft operational environment the defects and damage are likely to occur. Repair has to be carried out as per recommended and the requirement to replace back the sensors are important in order to monitor back the structure at post repair situation. The Lamb waves generated by using the PZT sensor can be used to monitor the surface structural integrity for damage or pristine condition. The effect of the lamb wave signals when surface condition of the aircraft component changes is concerned. One hundred data sets were recorded for the undamaged, damage and repair condition. An outlier analysis was used to analyze the situation by overlaying the isolated signal spectrum and the range of the voltage peak to peak (Vpp) mean values. Different signals were observed for different type of structural condition tested and more tests were required to make a conclusive solution.

A Fluid-Structure Interaction Model of Pressurised Composite Structures Subjected to Blast Loading

A numerical study into the dynamic behaviour of hybrid pressurised barrels manufactured from GLARE (Glass fibre Reinforced laminate) has been performed using the Arbitrary-Lagrangian-Eulerian (ALE) method that accounts for fluid structure interaction within the explicit finite element software RADIOSS. The results high-lighted the importance of the geometrical features of the closed barrel when assessing the shock wave propagation of the blast wave. Also the effect of pre-pressurisation was studied which proved significant in providing additional internal energy to the system. It was concluded that pressurisation should be accounted in all future studies to model the dynamic crack growth and structural integrity of typical aircraft structures subjected to blast.

The lead fatigue crack concept for aircraft structural integrity

Over many years of quantitative fractographic examination of fatigue cracking from in-service and full-scale fatigue tests of metallic airframe components, it has been consistently observed that the largest cracks formed have grown in an approximately exponential manner. These crack growth observations range from the initiation of cracks and their early growth from a few micrometers through to many millimetres in length. It appears that these lead cracks commence growing shortly after the airframe is introduced to the loading environment. Furthermore these cracks usually initiate from production-induced or, less frequently, inherent material discontinuities. Based on these two observations, an aircraft lifing methodology that is based on the results of fatigue testing programs utilising the lead crack concept has been developed and implemented as an additional tool in the determination of aircraft component fatigue lives in several Royal Australian Air Force (RAAF) fleet types. In this paper the lead crack concept is developed and its strengths and weaknesses are discussed. Examples of crack growth behaviour that are considered typical and representative of lead cracks are presented.

Effective Block Approach for Aircraft Damage Tolerance Analyses

A damage tolerance analysis method for aircraft structural integrity, designated as the effective block approach, has been developed and subsequently applied to predict the fatigue crack growth life of aircraft structures subjected to flight spectrum loading. The predictions by the method were made not based on constant-amplitude crack growth data, but instead relied on spectrum crack growth data obtained from previous full-scale fatigue tests or representative coupon fatigue tests. The spectrum crack growth data were normally measured from observations of the fracture surfaces by quantitative fractography. Verification and consistency studies of the method were performed against fatigue test results under different flight spectra. The predicted fatigue lives and inspection intervals for fracture-critical locations in aircraft structures were found to be in close agreement with representative coupon and full-scale fatigue tests. This study has demonstrated that the method provided significant advantages in damage tolerance analysis over conventional fatigue lifing approaches for military aircraft structures and components. It was also found that the work has improved the value of structural integrity advice provided for some Royal Australian Air Force air platforms and supports the safe operation of the fleets through to their planned withdrawal date.

Brazilian Air Force aircraft structural integrity program: an overview

This paper presents an overview of the activities developed by the Structural Integrity Group at the Institute of Aeronautics and Space - IAE, Brazil, as well as the status of ongoing work related to the life extension program for aircraft operated by the Brazilian Air Force BAF. The first BAF-operated airplane to undergo a DTA-based life extension was the F-5 fighter, in the mid 1990s. From 1998 to 2001, BAF worked on a life extension project for the BAF AT- 26 Xavante trainer. All analysis and tests were performed at IAE. The fatigue critical locations (FCLs) were presumed based upon structural design and maintenance data and also from exchange of technical information with other users of the airplane around the world. Following that work, BAF started in 2002 the extension of the operational life of the BAF T-25 “Universal”. The T-25 is the basic training airplane used by AFA - The Brazilian Air Force Academy. This airplane was also designed under the “safe-life” concept. As the T-25 fleet approached its service life limit, the Brazilian Air Force was questioning whether it could be kept in flight safely. The answer came through an extensive Damage Tolerance Analysis (DTA) program, briefly described in this paper. The current work on aircraft structural integrity is being performed for the BAF F-5 E/F that underwent an avionics and weapons system upgrade. Along with the increase in weight, new configurations and mission profiles were established. Again, a DTA program was proposed to be carried out in order to establish the reliability of the upgraded F-5 fleet. As a result of all the work described, the BAF has not reported any accident due to structural failure on aircraft submitted to Damage Tolerance Analysis.