Explore the vast range of titles available.

The environmental condition in which the Royal Malaysian Airforce is currently operating its aircraft is prone to corrosion. This is due to the high relative humidity and temperature. With most of its aircraft being in the legacy aircraft era, the aircraft’s main construction consists of the aluminium 2024 material. However, this material is prone to corrosion, thus reducing fatigue life and leading to fatigue failure. Using the concept of either Safe Life or Damage Tolerance as its fatigue design philosophy, the RMAF adopts the Aircraft Structure Integrity Program (ASIP) to monitor its structural integrity. With the current problem of not having the structural limitation on corrosion-damaged structure, the RMAF has embarked on its fatigue testing method. Finite Element (FE) studies and flight tests were conducted, and the outcome is summarized. The conclusion is that the longeron tested on the aircraft can withstand the operational load, and its yield strength is below the ultimate yield strength of the material. These research outcomes will also enhance the ASIP for other aircraft platforms in the RMAF fleet for its structure life assessment or service life extension program.

New experimental results on the effect of additional force impulse loading on the variation of the initial structure of the aircraft material (alloys D16, 2024-T3, VT22) at various stages of deformation are presented and a significant enhancement of its initial plasticity is achieved. Complex investigations into the material properties after a dynamic non-equilibrium process made it possible to describe the main regularities in the nature of deformation and fracture of materials, which allowed proposing general recommendations on using the revealed physical and mechanical regularities in the evaluation of strength of aircraft structures. First published online 10 May 2017

When designing thin-walled aircraft structures, as a rule, the main limitations are associated with ensuring stability. The objects of research in this work are composite z-shaped reinforced panels of the wing box of an aircraft. For the early stages of the design of load-bearing panels, it is necessary to evaluate the weight of the design decisions made, taking into account possible defects. The paper proposes an analytical technique for designing reinforced panels, taking into account the use of the condition of uniform stability and the presence of possible regulated defects in the skin that may occur in further operation. The task of determining the parameters of panels of minimum weight is reduced to minimizing the function of one variable, which is the ratio of the height to the reinforcement step. To take into account the indicated regulated defects, parametric studies were carried out using the finite element method, on the basis of which refinement coefficients were added to the analytical ratios of the stiffness parameters of the reinforced panels. The paper presents the results of studies of orthotropic panels with defects in the form of through holes under loading with compressive and tangential forces. The results are presented in the form of graphs that show possible changes in the critical stresses of thin skins with defects.

Aircraft icing remains one of the major problems in flight control and operation of aircraft. To avoid ice accretion various de-icing and anti-icing approaches have been developed. Recently, advanced studies were performed in the fields of numerical modeling of ice accretion, and its melting during an application of de-icing and anti-icing systems. However, little attention is still paid to the thermomechanical response of aircraft composite structures acting under icing conditions and using de-icing systems. The results of numerical simulations on thermomechanical response of a representative composite aircraft skin were analyzed, emphasizing representative scenarios of aircraft icing and acting of resistive heaters as a de-icing system. A detailed introduction to energy and mass transfer of ice accretion and melting phenomena was described to highlight additional major thermal effects. A comparative analysis between wet and dry air conditions was carried out. The simulations of aero-thermodynamic and structural coupling phenomena performed demonstrated mechanical behavior of composite aircraft structures under these complex conditions.

The calendar safety life of the surface protection system in aircraft structures is the time limit for it to be used without functional failure at a certain level of reliability and confidence. The reliability of such protection systems and the operational safety and economy of the structure are closely related. This paper firstly establishes two methods for determining the calendar safety life of aircraft structural protection systems under a single service environment and in multiple service environments. A method for determining the reliability of the calendar safety life of the aircraft structural protection system was proposed, and an expression of the relationship between the maintenance costs for the aircraft fleet and the reliability of the calendar safety life of the aircraft structural protection system based on the relationship between the amount of corrosion damage to the structural substrate and the corrosion time and the expression of the calendar safety life of the protection system was established. Finally, taking a hypothetical aircraft fuselage wall plate connection structure as an example, an alternating corrosion fatigue test with protection system specimens was carried out. The process for determining the calendar safety life of the structural protection system and its reliability are given. This method is important to ensure the safety of aircraft structures, improve the efficiency of use, and reduce maintenance costs. Generally speaking, the reliability of the calendar safety life of the structure is 99.9%, and after the analysis in this paper, the reliability of the structural protection system is about 70%.

Recent advances in aircraft materials and their manufacturing technologies have enabled progressive growth in innovative materials such as composites. Al-based, Mg-based, Ti-based alloys, ceramic-based, and polymer-based composites have been developed for the aerospace industry with outstanding properties. However, these materials still have some limitations such as insufficient mechanical properties, stress corrosion cracking, fretting wear, and corrosion. Subsequently, extensive studies have been conducted to develop aerospace materials that possess superior mechanical performance and are corrosion-resistant. Such materials can improve the performance as well as the life cycle cost. This review introduces the recent advancements in the development of composites for aircraft applications. Then it focuses on the studies conducted on composite materials developed for aircraft structures, followed by various fabrication techniques and then their applications in the aircraft industry. Finally, it summarizes the efforts made by the researchers so far and the challenges faced by them, followed by the future trends in aircraft materials.

Aircraft structures are highly susceptible to fatigue damage, particularly in thin-walled aluminum alloy components such as skin panels. Damage in the form of holes or material loss drastically reduces fatigue life and compromises structural safety, which makes effective repair strategies essential. This study presents an experimental investigation into the fatigue performance of EN AW-2024-T3 aluminum alloy plates with central openings subjected to uniform shear. Repair nodes were applied using two approaches: conventional riveted metal patches and adhesively bonded composite patches. Variants of patch geometry, thickness, and diameter were evaluated to determine their influence on load transfer, buckling response, and fatigue life. The results show that central holes significantly shorten fatigue life, with a 20 mm hole causing a 67% reduction and a 50 mm hole causing a 95% reduction when compared with undamaged plates. Riveted metal patches restored only part of the lost performance, as stress concentrators introduced by fastener holes initiated new fatigue cracks. In contrast, adhesively bonded composite patches provided a substantial improvement, extending fatigue life beyond that of the riveted solutions, improving buckling shape, and delaying crack initiation. Larger patches, particularly those combined with metallic inserts, proved most effective in restoring structural functionality. The findings confirm the effectiveness of bonded composite repairs as a lightweight and reliable method for extending fatigue life and enhancing the safety of damaged aircraft structures. The study highlights the importance of patch geometry and stiffness in the design of repair nodes.

Natural frequency is a critical dynamic parameter in designing the unmanned aerial vehicle (UAV) wing structure. The low-value natural frequency causes the aircraft wing to deflect easily while flying. Furthermore, the low-value natural frequency can causes resonance when this natural frequency is closest to the wind frequency. An increase in the UAV natural frequency can be conducted by increasing the wing structural stiffness. However, increased stiffness by enlarging dimensions is often followed by an increase in mass that is avoided in the design of aircraft structures. One technique that can be used to increase UAV natural frequency is by increasing the cross-sectional moment of inertia of the wing spar. In this study, an analysis of the effect of the wing spar cross-sectional inertia variations on the UAV natural frequency is conducted. Several wings spar cross-sectional profiles with different cross-sectional inertia are evaluated, such as U profile, circular, T profile, square, L profile, I profile, and Z profile. The results of the investigation showed that the wing spar using T profile is better than other profiles in increasing the wing structure natural frequency. This profile could increase the 1st elastic natural frequency by 22.9%, with only a 3.4 % increase in mass.

Including production considerations in the early design stages of aircraft structures is challenging. Production information is mostly known by experts and rarely formally documented such that it can be effectively used during the design process. Producibility is mostly considered after completing the design, resulting in increased cost and development time due to the late discovery of production issues. This paper presents a new model, called the Manufacturing Information Model (MIM), which supports the automatic inclusion of production considerations into the design process. The MIM provides a single source of truth and a generic structure to capture and organize production-related information in a product system. Furthermore, it provides compatibility analyses to automatically warn for or exclude infeasible designs. Analysis tools use the information stored within the MIM to calculate the mass, costs, and production rate of the product. To show the functionalities of the MIM, it has been applied to the conceptual design of a wing box at a Tier 1 company. This use case shows how the MIM supports trade-off decisions, as it allows for the identification of trends and the ranking of different manufacturing concepts. Overall, the MIM provides a structured and formal approach to include production information in the conceptual design, improving the decision-making process.

Flutter is a self‐excited vibration under the interaction of the inertial force, aerodynamic force, and elastic force of the structure. After the flutter occurs, the aircraft structures will exhibit limit cycle oscillation, which will cause catastrophic accidents or fatigue damage to the structures. Therefore, it is of great theoretical and practical significance to study the aeroelastic characteristics and flutter control for improving the aeroelastic stability of aircraft structures. This paper reviews the recent advances in aeroelastic analysis and flutter control of wings and panel structures. The mechanism of aeroelastic flutter of wings and panels is presented. The research methods of aeroelastic flutter for different structures developed in recent years are briefly summarized. Various control strategies including the linear and nonlinear control algorithms as well as the active flutter control results of wings and panels are presented. Finally, the paper ends with conclusions, which highlight challenges of the development in aeroelastic analysis and flutter control, and provide a brief outlook on the future investigations. This study aims to present a comprehensive understanding of aeroelastic analysis and flutter control. It can also provide guidance on the design of new wings and panel structures for improving their aeroelastic stability.