Sabih Ahmad Khan

This study aims to evaluate Remaining Useful Life (RUL) of supersonic aircraft using a two-fold integrated approach that combines experimental and computational methods. A comprehensive framework was developed that included the identification of Fatigue Critical Locations (FCLs), aircraft instrumentation, Ground Strain Survey (GSS), in-flight testing, and RUL estimation. Internal and external CAD models of the aircraft were developed. CFD analysis on various critical flight maneuvers inside the flight envelope was performed using an external CAD model. Subsequently, Finite Element Analysis (FEA) was carried out under severe CFD load cases. Based on FEA results, five FCLs were identified on the wing, two on the vertical tail, and two on the horizontal stabilizer. In the experimental phase, the aircraft was instrumented with strain gauges, accelerometers, and a Data Acquisition Unit (DAQ) to collect in-flight structural testing data. The FEA model was calibrated through GSS. The calibrated weights were applied to the aircraft wings; corresponding experimental strains were recorded and validated against the FE model, yielding a percentage error of 5%–11%. Five flight tests, carefully selected based on historical aircraft trends and critical aircraft maneuvers, were performed to obtain in-flight strain data. The Palmgren-Miner rule was used to evaluate damage caused by each flight test against each structural member on identified FCLs. The total damage was estimated by extending the assessment over the aircraft’s operational life span, surpassing 3000 OEM-defined flying hours. The Main Spar of the wing was found to have the least life of 7089.52 AFHRs, making it the most critical structural member of the aircraft. This framework provides a systematic methodology for estimating RUL, monitoring structural health, addressing gaps in existing methods for evaluating catastrophic failure, and serving as a valuable source for life extension studies in the aviation sector.

This research paper presents an application of the integrated process and product design (IPPD) approach for selecting the best joint configuration for dissimilar material joining in the early product design phase. The proposed methodology integrates the multi-criteria decision making (MCDM) approach with quality function deployment (QFD) to identify the key criteria for joint selection, including load-carrying capacity, size, cost per joint, ease of manufacturing, time consumption and deformation. Three types of joints (rivet, weld and adhesive) and two hybrid joints (adhesive-weld and adhesive-rivet) are considered for three dissimilar material configurations: carbon fiber-reinforced plastic (CFRP) aluminum, CFRP steel, and aluminum-steel. QFD is utilised to transform job requirements into design criteria, and in the second phase, the Technique for Order Preference by Similarity to Ideal Solution (TOPSIS) is employed to choose the optimal joint configuration based on the weighted criteria acquired in the previous phase. The selected joint configuration is then validated through experimental study. The findings demonstrate that the proposed IPPD approach with QFD-TOPSIS techniques is highly effective for selecting mechanical joints for dissimilar material joining in the early design phase. The study concludes that the adhesive-rivet hybrid joint is the optimal solution among all alternatives. The proposed methodology can ultimately lead to improved product reliability and performance, as well as reduced development time and cost.

The determination of suitable testing and qualification procedures for fiber-reinforced polymer matrix composite structures is an active area of research due to the increased demand, especially in the field of aerospace. This research illustrates the development of a generic qualification framework for a composite-based main landing gear strut of a lightweight aircraft. For this purpose, a landing gear strut composed of T700 carbon fiber/epoxy material was designed and analyzed for a given lightweight aircraft having mass of 1600 kg. Computational analysis was performed on ABAQUS CAE^® to evaluate the maximum stresses and critical failure modes encountered during one-point landing condition as defined in the UAV Systems Airworthiness Requirements (USAR) and Air Worthiness Standards FAA FAR Part 23. A three-step qualification framework including material, process and product-based qualification was then proposed against these maximum stresses and failure modes. The proposed framework revolves around the destructive testing of specimens initially as per ASTM standards D 7264 and D 2344, followed by defining the autoclave process parameters and customized testing of thick specimens to evaluate material strength against the maximum stresses in specific failure modes of the main landing gear strut. Once the desired strength of the specimens was achieved based on material and process qualifications, qualification criteria for the main landing gear strut were proposed which would not only serve as an alternative to drop test the landing gear struts as defined in air worthiness standards during mass production, but would also give confidence to manufacturers to undertake the manufacturing of main landing gear struts using qualified material and process parameters.

The design and development of high-strength and low-weight composite landing gear struts is still a challenge in today’s world. In this study, a selection methodology for fiber-reinforced composite material for retractable main landing gear struts for specified lightweight aircraft up to 1600 kg mass is proposed. Four different fiber-reinforced composite materials, two each from the glass-fiber and carbon-fiber families, including E-glass fiber/epoxy, S-glass fiber/epoxy, T300 carbon fiber/epoxy, and AS carbon fiber/epoxy, were considered for analysis. For the design and analysis of a main landing gear strut, maximum landing loads for one point and two point landing conditions were calculated using FAA FAR 23 airworthiness requirements. Materials were categorized based on their strength-to-weight ratio and the Tsai-Wu failure criterion. Landing gear struts meeting the Tsai-Wu failure criterion, and having a maximum strength-to-weight ratio, were then modeled for performance under a collision detection test. This research concludes that T300 carbon fibre/epoxy is a recommended material for the manufacture of landing gear struts for specified lightweight aircraft.