Damage Tolerance Requirements Research Articles

Damage Tolerance Behaviour of Stiffened Crown Panel of a Transport Aircraft Fuselage

Ever since the introduction of damage tolerance requirements in the aviation regulations, efforts continue to be made to prevent catastrophic failures due to damages present in the structure. It has also been realized that damage detection is the weakest link in the whole process of damage tolerance design to maintain continued airworthiness. The major components of the aircraft structure consist of both integral and riveted panels of sheets and stringers which are employed in fuselage skin panels, spar webs and stiffeners. In spite of all precautions, cracks or damages may arise in many of these primary structural members. These cracks cause stiffness degradation and reduce the total load-carrying capacity of the structure. In this paper, the damage tolerance behaviour of fuselage crown panel both integral and riveted stiffened panel configurations of Aluminium alloy 2024-T351 are studied using finite element based tools using crack growth analysis methods. The crack growth behaviour of both integral and riveted stiffened panels of aircraft fuselage having same geometrical configuration and subjected to uniformly distributed tensile loads is investigated. For this, a metallic stiffened panel with eight stringers, representative of crown panel of a transport aircraft fuselage is analysed with a centre skin crack propagating through the stringers. Stress intensity factors and fatigue crack propagation rates at the progressive crack tip of both types of the stiffened panels are computed by using Modified Virtual Crack Closure Integral (MVCCI) method. The stiffened panels fatigue crack growth rate was computed by using Paris law under constant amplitude fatigue loads. The analysis results show that integral stiffened panel causes higher stress intensity factor and less load bearing capability than riveted stiffened panel which has better damage tolerant capability in comparison to the integrally stiffened panel.

Conceptual Definition of A New Very-Light-Jet Aircraft Configuration

Israel Aircraft Industries (IAI) is a world leader in the design and manufacturing of business aircraft. IAI has developed and introduced to the market several biz-jets (Westwind I and II, Astra-AKA Gulfstream G-100, and Galaxy- AKA Gulfstream G-200). All these aircraft have been in the midsize and recently "super-midsize" category. During the last two years, IAI has been evaluating a new aircraft design, the ProJet that belongs to the new, emerging category of "very-light jets" (VLJ). This lightweight twin turbofan aircraft is characterized by its small size and low target price, compared to existing IAI designs. The ProJet will be certificated to FAR23 normal category requirements, including single pilot operation. The ProJet exploits newly developed technologies and capabilities, such as: new, small and efficient turbofan engines; advanced low cost and sophisticated avionic systems; advanced low cost mechanical and electrical systems; rapid and efficient development and production methods. All these attributes will provide the capability of introducing a very low cost aircraft into the emerging new VLJ market sector. The ProJet is a twin turbofan, low wing, T-tail airplane utilizing an all-metal airframe; designed in accordance with damage tolerance requirements. It features a comfortable pressurized cabin with optimized seating for 6 occupants, including 2 pilots, and excellent performance. The ProJet will have a VFR range of 1200nm, a cruise ceiling of 41,000ft, and a cruise speed of 365kt. It will take-off and land on runways shorter than 3000ft. Generous allowance for baggage stowage is provided in the pressurized internal baggage compartment, and in the external rear baggage compartment. The ProJet flight deck features an all-glass cockpit with an integrated avionics package, designed for ease of operation and reduced workload. Composite materials are utilized in flight control surfaces and secondary structures. Two FADEC controlled turbofan engines mounted on the upper aft fuselage supply the ProJet power. Each engine generates about 1300 pounds static installed thrust at ISA +10°C.

Design and Testing of Impacted Stiffened CFRP Panels under Compression with the VERTEX Test Rig

Aeronautical composite primary structures must evidence sufficient residual strength in the presence of damage for compliance with damage tolerance requirements. The study of stiffener debonding on panels subjected to compression after impact is performed in that scope. Compression leads to the buckling of the skin between the stiffeners, and thus a complex loading of the bonding between the skin and the stiffener. This paper describes the development of a stiffened specimen for the VERTEX multiaxial test rig as a first step towards the study of the damage tolerance evaluation of stiffened structures, under combined loadings and at the intermediate scale of the test pyramid. By using virtual testing, the specimen was designed to produce the phenomenology of interest as the first damage, i.e., the debonding of the stiffener from the centre. Three samples were manufactured and subjected to low velocity impacts at various locations and energies. Then the three samples were subjected to compression after impact, up to the stiffener debonding, under a post-buckling regime of the skin. Test loading evolution is described with force fluxes and global strains, obtained from in situ stereo-correlation. The different impacts were found to give different types of damage but similar residual strength to compression after impact.

Composites sandwich structures certification: A case study

The motivation of this work is to show a case study represented by a test article in tapered sandwich structure with a mid-ramp region in order to reduce certification cost of the full-scale structure. Also, this work provides a compelling case for the standardization of Load Enhancement Factor for certification purposes. Therefore, the development of this work is based on the definition of the test article with a manufacturing defect and impact damage, for compliance static and damage tolerance requirements with damage no-growth approach. The results obtained showed a good strategy to reduce risk and cost of the certification development.

A study of the damage tolerance of composite-metal hybrid joints reinforced by multiple and penetrative thin pins

The application of adhesive bonding technology in aircraft structures can reduce the total wight greatly, but the bonded joints are very sensitive to the possible manufacturing defects and damages during service operations, which makes them difficult to meet the damage tolerance requirements of the current transport airplane structures. In this study, the damage tolerance of composite-metal hybrid joints reinforced by multiple and penetrative thin pins was studied. The damage tolerance performance of the composite-metal joint is supposed to be enhanced by multiple through-the-thickness penetrative thin reinforcements in the bonding region, and the thin reinforcements were bonded together with both the composite and metallic joint plates. Both experimental tests and finite element simulations were conducted to investigate the effects of the through-the-thickness reinforcements on the damage tolerance performance of the joints with and without pre-fabricated disbond defects. Through the comparative analyses, it was found that the penetrative thin pins in the bonding region significantly improved the static load carrying capacity, the failure strain, the fracture energy, and the fatigue lives of the composite-metal bonded joints. Moreover, the reinforcements decreased the sensitivity of the bonded joints to the disbond defects in the bonding region. The damage tolerance performance of the composite-metal adhesively bonded joints was significantly increased by the through-the-thickness penetrative reinforcements and the enhancement mechanism was revealed by the combined analysis of test results and simulation results.

Structural Static/Fatigue/Damage Tolerance Comprehensive Design Optimization of Composite Laminate

This paper develops a structural static/fatigue/damage tolerance comprehensive design method for composite laminated structures under mechanical–thermal coupling environments. The developed design method can finely quantify the additional stress and deformation caused by the mechanical–thermal coupling effect. In addition, a new damage evolution-based fatigue reliability evaluation method is proposed, and the structural fatigue behavior can be described and predicted. The different failure modes of composite material and their correlations are considered during the damage evolution analysis. The structural possible accidental damage can also be taken into account by applying the global preset damage in design stage. Thus, the static requirement, fatigue requirement, and damage tolerance requirement can all be considered simultaneously during the design process. The structural comprehensive design of composite laminate is conducted by virtue of the two-level optimization method. A case of a notched composite laminate is employed to verify the proposed method. In addition, a case of a civil aircraft wing structure is used to explain the feasibility of the developed method in dealing with structural design problems in engineering.

Further Studies into Crack Growth in Additively Manufactured Materials.

Understanding and characterizing crack growth is central to meeting the damage tolerance and durability requirements delineated in USAF Structures Bulletin EZ-SB-19-01 for the utilization of additive manufacturing (AM) in the sustainment of aging aircraft. In this context, the present paper discusses the effect of different AM processes, different build directions, and the variability in the crack growth rates related to AM Ti-6Al-4V, AM Inconel 625, and AM 17-4 PH stainless steel. This study reveals that crack growth in these three AM materials can be captured using the Hartman–Schijve crack growth equation and that the variability in the various da/dN versus ΔK curves can be modeled by allowing the terms ΔKthr and A to vary. It is also shown that for the AM Ti-6AL-4V processes considered, the variability in the cyclic fracture toughness appears to be greatest for specimens manufactured using selective layer melting (SLM).

Integrity Management of Offshore Structures With Emphasis on Design for Structural Damage Tolerance

AbstractBased on relevant accident experiences with oil and gas platforms, structural integrity management of offshore structures is briefly outlined, including adequate design criteria, fabrication and operational procedures, as well as life cycle quality assurance and control. The focus is on developing an operational design standard for accidental collapse limit states to ensure robustness or damage tolerance. The focus is to ensure an acceptable safety level against progressive failure leading to total loss in view of initial damage caused by accidental actions due to operational errors and abnormal structural damage due to fabrication errors and abnormal deterioration during operation as well as the actions on the damaged structure and inherent uncertainties. Moreover, the damage tolerance required for achieving safety by inspection, monitoring and repair strategies, is briefly addressed. While the basic damage tolerance requirement refers to the survival of the structure in certain damage conditions, wider aspects of robustness in terms of the structure’s sensitivity to the deviation of action effects and resistances from normal conditions are also briefly addressed. In particular, it is suggested to provide robustness in cases where the structural performance is sensitive to uncertain parameters, by choosing conservative values of these parameters.

Digital twin approach for damage-tolerant mission planning under uncertainty

The digital twin paradigm that integrates the information obtained from sensor data, physics models, as well as operational and inspection/maintenance/repair history of a system (or a component) of interest, can potentially be used to optimize operational parameters of the system in order to achieve a desired performance or reliability goal. In this article, we develop a methodology for intelligent mission planning using the digital twin approach, with the objective of performing the required work while meeting the damage tolerance requirement. The proposed approach has three components: damage diagnosis, damage prognosis, and mission optimization. All three components are affected by uncertainty regarding system properties, operational parameters, loading and environment, as well as uncertainties in sensor data and prediction models. Therefore the proposed methodology includes the quantification of the uncertainty in diagnosis, prognosis, and optimization, considering both aleatory and epistemic uncertainty sources. We discuss an illustrative fatigue crack growth experiment to demonstrate the methodology for a simple mechanical component, and build a digital twin for the component. Using a laboratory experiment that utilizes the digital twin, we show how the trio of probabilistic diagnosis, prognosis, and mission planning can be used in conjunction with the digital twin of the component of interest to optimize the crack growth over single or multiple missions of fatigue loading, thus optimizing the interval between successive inspection, maintenance, and repair actions.

Bio-inspired design for enhanced damage tolerance of self-reinforced polypropylene/carbon fibre polypropylene hybrid composites

In this work, we investigate the toughness of an inter-layer Self-Reinforced Polypropylene/Carbon Fibre Polypropylene (SRPP/CFPP) cross-ply hybrid composite and devise strategies to improve two aspects of its damage tolerance: (i) increasing the energy dissipation capability and (ii) enhancing the impact damage tolerance. To this end, we introduced discontinuities in the form of laser-cuts across the fibres of the CFPP plies, tailoring two patterns of laser-cuts to meet each specific damage tolerance requirement. We conducted penetration impact and Double Edge Notched Tensile (DEN-T) tests. The DEN-T tests, analysed via the Essential Work of Fracture method, show that engineering the microstructure successfully diffused damage. This resulted in a great increase in energy dissipation capability — 90% higher than a reference non-engineered structure. Engineering the microstructure of impact samples has led to enhanced impact damage tolerance with increased energy dissipation at a sub-critical level and delayed critical failure.

Optimal design of composite lateral wing upper covers. Part II: Nonlinear buckling analysis

The present investigation is devoted to a development of new optimal design concepts of aircraft lateral wing upper covers made of advanced composite materials. In the second part a stiffened composite panel with the best weight/design performance obtained from the linear buckling analysis (Part I) is verified by the nonlinear buckling analysis and re-optimized in the case of necessity. Additionally an effect of shear and fuel pressure as well as an effect of skin post-buckling on its performance is investigated. Three rib bays laminated composite panels with HAT-stiffeners were modeled with ANSYS finite element code to study their buckling behavior as a function of skin and stiffener lay-ups, stiffener height, stiffener top and root width. Due to the large dimension of numerical problems to be solved, an optimization methodology was developed employing the method of experimental design and response surface technique. Weight optimization problems were solved for four load levels, taking into account manufacturing, repairability and damage tolerance requirements. Optimal results were verified successfully using ANSYS shared-node model.

Weight optimal design of lateral wing upper covers made of composite materials

The present investigation is devoted to the development of a new optimal design of lateral wing upper covers made of advanced composite materials, with special emphasis on closer conformity of the developed finite element analysis and operational requirements for aircraft wing panels. In the first stage, 24 weight optimization problems based on linear buckling analysis were solved for the laminated composite panels with three types of stiffener, two stiffener pitches and four load levels, taking into account manufacturing, reparability and damage tolerance requirements. In the second stage, a composite panel with the best weight/design performance from the previous study was verified by nonlinear buckling analysis and optimization to investigate the effect of shear and fuel pressure on the performance of stiffened panels, and their behaviour under skin post-buckling. Three rib-bay laminated composite panels with T-, I- and HAT-stiffeners were modelled with ANSYS, NASTRAN and ABAQUS finite element codes to study their buckling behaviour as a function of skin and stiffener lay-ups, stiffener height, stiffener top and root width. Owing to the large dimension of numerical problems to be solved, an optimization methodology was developed employing the method of experimental design and response surface technique. Optimal results obtained in terms of cross-sectional areas were verified successfully using ANSYS and ABAQUS shared-node models and a NASTRAN rigid-linked model, and were used later to estimate the weight of the Advanced Low Cost Aircraft Structures (ALCAS) lateral wing upper cover.

Optimal design of composite lateral wing upper covers. Part I: Linear buckling analysis

The present investigation is devoted to a development of new optimal design concepts of aircraft lateral wing upper covers made of advanced composite materials. In the first part three rib bays laminated composite panels with T, I and HAT-stiffeners were modeled with ANSYS and NASTRAN finite element codes to investigate their buckling behavior as a function of skin and stiffener lay-ups, stiffener height, stiffener top and root width. Due to the large dimension of numerical problems and large number of optimization tasks to be solved, an optimization methodology was developed employing the method of experimental design, response surface technique and linear buckling analysis. Weight optimization problems were solved for the laminated composite panels with three types of stiffeners, two stiffener pitches and four load levels taking into account manufacturing, repairability and damage tolerance requirements. The composite panel with the best stiffener type was identified for the subsequent nonlinear buckling analysis and optimization presented in the Part II. Optimal results were verified successfully using ANSYS shared-node and NASTRAN rigid-linked models.

Fatigue Crack Retardation in LSP and Sp Treated Aluminium Specimens

Highly loaded aircraft components have to fulfill strict fatigue and damage tolerance requirements. For some components besides the crack initiation mainly the fatigue crack propagation behavior is the main design criteria. To improve the crack propagation behavior of a component several methods are known or have been described in literature. For thin aircraft panels i.e. the application of crenellations [1] or bonded doublers [2, 3] can be a solution. For thick structures mainly the introduction of compressive residual stresses is beneficial. In this paper the potential of compressive residual stresses obtained by Laser Shock Peening (LSP) and Shot Peening (SP) is investigated. By means of Laser Shock Peening the residual compressive stress field can extend much deeper below the treated surface than that produced by conventional Shot Peening (i.e. with steel or ceramic balls) [4, 5]. The effect of such deep compressive stress profile results in a significantly higher benefit in fatigue behavior after Laser Shock Peening or after the combination of Laser Shock Peening and Shot Peening on top. The measurement of residual stresses as a depth profile has been performed by incremental hole drilling (ICHD) and contour method. Finally crack propagation tests have been carried out to validate the process technology approach.

Damage Tolerance of Pre-Stressed Composite Panels Under Impact Loads

An experimental test campaign studied the structural integrity of carbon fibre/epoxy panels preloaded in tension or compression then subjected to gas gun impact tests causing significant damage. The test programme used representative composite aircraft fuselage panels composed of aerospace carbon fibre toughened epoxy prepreg laminates. Preload levels in tension were representative of design limit loads for fuselage panels of this size, and maximum compression preloads were in the post-buckle region. Two main impact scenarios were considered: notch damage from a 12 mm steel cube projectile, at velocities in the range 93–136 m/s; blunt impact damage from 25 mm diameter glass balls, at velocities 64–86 m/s. The combined influence of preload and impact damage on panel residual strengths was measured and results analysed in the context of damage tolerance requirements for composite aircraft panels. The tests showed structural integrity well above design limit loads for composite panels preloaded in tension and compression with visible notch impact damage from hard body impact tests. However, blunt impact tests on buckled compression loaded panels caused large delamination damage regions which lowered plate bending stiffness and reduced significantly compression strengths in buckling.

Optimal design of composite upper covers of lateral wings with the effect of rib attachment to stiffener webs

The present investigation is devoted to the development of new optimal design concepts that exploit the full potential of advanced composite materials in the upper covers of aircraft lateral wings. A finite-element simulation of three-rib-bay laminated composite panels with T-stiffeners and a stiffener pitch of 200 mm is carried out using ANSYS to investigate the effect of rib attachment to stiffener webs on the performance of stiffened panels in terms of their buckling behavior and in relation to skin and stiffener lay-ups, stiffener height, and root width. Due to the large dimension of numerical problems to be solved, an optimization methodology is developed employing the method of experimental design and the response surface technique. Minimal-weight optimization problems were solved for four load levels with account of manufacturing, repairability, and damage tolerance requirements. The optimal results were verified successfully by using the ANSYS and ABAQUS shared-node models.

Damage Tolerance and Reliability Assessment under Random Markovian Loads

Accounting for load uncertainties plays an important role in the design of safe structural components of aircrafts under damage tolerance requirements. The purpose of this paper is to develop a reliability assessment technique for cracked structures submitted to non-stationary random fatigue loads modeled by first-order Markov chains with discrete state space and identified from in-flight measurements. The strategy based on a multi-level version of the cross-entropy method consists in progressively updating the tran- sition probability matrix in order to generate load sequences of increasing severity which are likely to cause failure. The proposed method is applied to a cracked M(T) specimen under the defined random fatigue loads. Load cycle interactions and retardation effects are accounted for by means of the PREFFAS crack closure model. The efficiency of the proposed approach in terms of computational cost is clearly observed for rare failure events in comparison with direct Monte Carlo simulations. In addition to the failure probability estimate, the multi-level cross-entropy method provides the analyst with information on the most probable load sequences at failure.

Design of a Motorcycle Composite Swing-Arm by Means of Multi-objective Optimisation

A study for the replacement of a metallic swing-arm of a high performance motorcycle with a composite part is presented. Considering the high structural effectiveness of the original metallic component, the case study evaluates the potential of composites in a challenging application. The FE model of the original component is developed to evaluate the structural performance in the most significant load conditions. A manufacturing process, based on a RTM technique, is proposed and analysed in order to develop realistic design hypotheses. The design approach is based on an optimisation process with 60 design variables. A constrained multi-objective genetic algorithm is applied to identify the solutions representing the best trade-off between mass reduction and improvement of torsional stiffness. Results show that composite materials can enhance the structural efficiency of the original metallic part, even considering technological limitations and damage tolerance requirements.

Compression After Impact Testing of Sandwich Composites for Usage on Expendable Launch Vehicles

Composite material usage is necessary on NASA’s future launch vehicles in order to obtain a low mass vehicle. While aircraft and launch vehicles that utilize load-bearing composite components have many similar damage tolerance requirements, the distinct differences between a part that has a lifetime of ∼500 s (one launch) and can be inspected in detail before use and one that has a lifetime of many tens of thousands of flight hours and can only undergo a ‘walk around’ inspection before each flight (commercial transport) needs to be taken into account. This article presents these differences and uses data from the ARES I composite interstage as an example of how to arrive at preliminary compression after impact strength values for the sandwich structure in the acerage of this part using residual strength curves. Results show that if severity of damage can be quantified by a nondestructive method (other than dent depth), the mass of the structure can be reduced due to better characterization of the damage.

Damage tolerance demonstration testing for the Australian F/A-18

Over recent years Australia has been involved in a number of full-scale fatigue testing programs in support of the through-life structural integrity of the Royal Australian Air Force’s (RAAF) F/A-18 fleet. It was recognised early in the acquisition cycle that the certification testing conducted by the manufacturer failed to considered damage tolerance requirements and would be unlikely to cover the typically more severe and diverse RAAF operations. Given similar aircraft structural integrity management philosophies, major benefits were to be realised through collaboration with the Canadian Forces (CF). In particular, as fatigue testing under representative CF/RAAF loading was the basis for both countries’ structural integrity management, the International Follow-On Structural Test Project (IFOSTP) was successfully concluded. This paper emphasises the Australian components of IFOSTP, including the damage tolerance testing and demonstration for the aft fuselage that incorporated the simultaneous application of both manoeuvre and dynamic buffet loads. Many of the innovations and consequences of this work program are highlighted, and may be applicable to future fighter aircraft structural integrity programs.