Gas turbine maintenance strategy relies heavily on accurate estimation of critical component life consumption of gas turbine engines during their operations. The equivalent operating hours (EOH) is a useful concept to measure the engine life consumption and support condition-based maintenance planning for gas turbine engines and their critical components. However, the current EOH calculation methods are mostly empirical and engine-specific, relying on vast operating data and experience. This paper introduces a novel physics-based method to estimate the EOH of the high-pressure turbine rotor blades of a gas turbine engine based on the damages caused by creep and low-cycle fatigue (creep-LCF) interactions. The method has been applied to a typical turbofan engine taking both 440-minute long-haul flight at one flight per day and 60-minute short-haul flight at two flights per day. A comparison of the predicted damages and life consumptions indicates that the creep EOH and also the creep damage of the engine of the short-haul aircraft is about 1.38 times that of the engine of the long-haul aircraft, the LCF equivalent operating cycles (EOC) and also the LCF damage of the engine of the short-haul aircraft is about 2.0 times that of the engine of the long-haul aircraft, and the total damages are more affected by the creep damage than the LCF damage with the creep damage being 6.78 times the LCF damage for the engine of the short-haul aircraft and 9.81 times for the engine of the long-haul aircraft. In addition, the total EOH or the total damage of the engine of the short-haul aircraft is about 1.44 times that of the engine of the long-haul aircraft. The proposed method shows a great potential to provide a quick estimate of the life consumption of gas turbine engines for condition monitoring, and it can be applied to other types of gas turbine engines.

Standard quadrotors exhibit limited mobility due to inherent underactuation: they only have four independent control inputs, whereas their position and attitude in space are defined by six degrees of freedom (DOF). Consequently, a quadrotor’s pose cannot track an arbitrary trajectory over time. To address this limitation, a novel actuation concept has been proposed, wherein the quadrotor’s propellers can tilt around their axes relative to the main body–forming a vector quadrotor. To achieve more accurate trajectory tracking tailored to the specific characteristics of this vector quadrotor model, we propose a novel control strategy. First, we integrate special orthogonal group SO(3) theory with model compensation control: SO(3) theory enables accurate modeling of the aircraft’s rotational dynamics, while model compensation control mitigates unmodeled dynamics and external disturbances, thereby ensuring robustness across diverse operating conditions. Second, we introduce the sequential quadratic programming (SQP) method for control allocation; this method not only enables efficient computation of control inputs but also optimises the allocation of control resources, which enhances system performance–particularly in complex manoeuvering scenarios. Finally, we integrate the SO(3)-based controller with the SQP-based control allocation module to form a unified control system. The effectiveness of this proposed approach is validated via simulation results. These results demonstrate improved trajectory tracking accuracy and enhanced robustness against disturbances, thus confirming the potential of our method for practical applications.

Flow separation in highly loaded axial compressors remains a major barrier to performance, motivating the search for active flow control strategies. This study investigates air injection to energise low-momentum endwall flow in a tandem stator configuration, representing the first investigation of its kind for tandem vanes. A numerical investigation was conducted, starting with a smooth-casing reference case and progressing to parametric studies of slot geometry (inclination $\alpha $, jet angle $\beta $, radius of curvature ${R_c}$, circumferential width ${w_c}$), relative injection mass flow rate ${\dot m_{inj}}/{\dot m_{stall}}$ and axial location $\zeta $. The results show how each parameter influences efficiency and pressure ratio, yielding design guidelines: shallow $\alpha $, moderate $\beta $ towards the separation zone, relatively large ${R_c}$ and a balanced ${w_c}$–${\dot m_{inj}}/{\dot m_{stall}}$ combination, best captured through the momentum coefficient ${C_u}$ and velocity ratio ${u_{inj}}/{u_\infty }$. Injection near $\zeta \approx 1.2$ (just upstream of separation) proved most effective, and off-design simulations showed larger efficiency gains towards de-throttled conditions, although stall margin was unaffected. Robustness was confirmed through turbulence-model comparisons and injector turbulence variations, which consistently reproduced suppression of suction-side separation. An integrated analysis of aerodynamic losses further showed that injection strategies remain beneficial when loss penalties are considered. The study thus establishes transferable guidelines for injector design in tandem stators, providing a foundation for future optimisation and experimental validation.

This paper presents a comprehensive experimental investigation into the shock characteristics associated with a low-thrust, low-shock separation mechanism incorporating Mild Detonating Cord (MDC) within a rubber bellow interface. Two test configurations were developed with varying explosive charge masses to study their influence on pressure generation and shock propagation. Linear accelerometers and high-speed pressure transducers were employed to capture transient dynamic responses at both piston and cylinder interfaces. The results demonstrate a significant reduction in peak pressure and shock levels, especially in the second test configuration, where the explosive mass was reduced to 60% of the initial configuration. The shock response spectrum (SRS) analysis confirms that the lower charge mass leads to proportionally reduced shock amplitudes across the frequency range of interest. Furthermore, comparative assessment of shock levels reveals a significant reduction of shock levels as compared to conventional separation mechanisms, such as a flexible linear-shaped explosive charge (FLSC) mass or a separation bolt actuated with a pyro cartridge. The experimental pressure values are shown to correlate well with theoretical predictions, validating the design approach. These findings provide critical insights into tailoring explosive-based separation mechanisms for sensitive payload environments, highlighting the importance of confined detonation and charge optimisation in mitigating pyroshock.

The effect of the bio-inspired leading-edge modifications on the aerodynamic performance of non-slender delta wing models was investigated in a low-speed wind tunnel using force and surface pressure measurements. The measurements were performed at a Reynolds number of $Re = 1 \times {10^5}$ over an angle-of-attack range from $ - 4^\circ $ to $30^\circ $. Seven different sharp-edged delta wing models with a 45-degree sweep angle (${\rm{\varLambda }}$), including a base wing, were used to study the effect of sinusoidal and saw-tooth leading-edge modifications. Sinusoidal leading-edge wing designs were inspired by the leading-edge tubercles of the humpback whale’s pectoral fins. The results indicate that the bio-inspired wing modifications resulted in a delay in the stall angle by 4 degrees, smoother stall characteristics, a higher maximum lift coefficient, and increased post-stall lift. The drag coefficient of the modified wings was observed as higher than that of the base wing model. Regarding the longitudinal static stability, leading-edge modifications decreased the stability of the wing as the angle-of-attack surpassed $\alpha = 17^\circ $.

In the field of parafoil airdrop path planning, the inherent complexity and time-sensitive nature of mission requirements necessitate rapid path generation through low-order mathematical models that approximate the system’s true dynamics. This study presents a novel sparse identification framework for constructing a parafoil path planning approximate model. Leveraging high-fidelity 9-degree-of-freedom (9 DOF) dynamic simulation data as training inputs, our method identifies simple nonlinear relationships between 3D positional coordinates (for spatial targeting) and yaw angle (for directional control), which are critical path planning parameters. Compared to conventional 4 DOF models, experimental validation using field airdrop data reveals that the proposed sparse model achieves enhanced predictive accuracy while maintaining computational efficiency. Quantitative analysis demonstrates reductions in root mean square error (RMSE) by approximately 12.96% (horizontal position), 54.44% (height) and 37.96% (yaw angle). The efficacy is further confirmed through successful fixed-point homing across diverse initial deployment scenarios, underscoring its potential for parafoil path planning.

In the process of utilising machine vision-assisted large aircraft component docking assembly, due to the occlusion induced by process equipment such as assembly tooling, the features on the calibration board cannot be extracted by each camera at the same time, resulting in calibration difficulties or calibration failure. This paper aims to propose a stereo calibration method for multi-cameras in large aircraft component assembly to improve calibration accuracy. Firstly, the sub-pixel edge extraction method based on Canny-Zernike is proposed to accurately extract the circular edges and circle centres of the calibration board, and the Zernike moment model is improved. The circle centre sorting method based on the triangular markers is introduced to realise the sorting of circle centres on the calibration board. Secondly, the intrinsic and extrinsic parameter models of multi-cameras and the visual parameter models between cameras are constructed, and Zhang’s calibration method and indirect calibration method are integrated to solve the parameters. Subsequently, the distortion correction model is optimised by Levenberg-Marquardt. Finally, experiments are performed to test the proposed method. The results show that the proposed method, compared with uncalibration and Zhang’s calibration method, the proposed method achieves stereo calibration of the multi-cameras under complex working conditions, enhances the calibration accuracy and improves the quality of the large aircraft component docking assembly.

Collaborative robots (cobots) have emerged as a pivotal paradigm for the upcoming leap to Industry 5.0. In recent years, the range of applications has expanded significantly, particularly in assembly tasks within the manufacturing industry. The primary goal of this paper is to review the application of cobots in industrial assembly tasks, highlighting possibilities for innovative research in smart robotics, including prospects for challenging applications in aircraft final assembly processes. The paper systematically reviews recent literature to analyse the use of collaborative robotics in industrial assembly tasks, encompassing characterisation of application environments, motivations, characteristics and outcomes of relevant use cases across various industrial segments. Additionally, it reviews a set of innovative technological patents issued by the aeronautical industry over the past 14 years, highlighting trending projects in industry. The investigation reveals that the automotive and electronics industries remain at the forefront of cobot applications, mainly for tasks like pick-and-place operations and component manipulation. Applications in open work cells, where humans and robots operate at supportive or sequential interaction levels, using conventional communication interfaces and camera-assisted technologies, have been the most prevalent. The review identifies potential opportunities and key aspects from future application scenarios for cobots. The findings are relevant to the industrial robotics community, emphasising the need for novel applied research on human–robot colla boration in aeronautical industry.

The present work investigates the thermochemical non-equilibrium effect in the DLR combustor using a two-temperature model combined with vibration-chemistry coupling model. Two operating conditions with inflow Mach 2 and 6 are selected for study. The simulation results illustrate that translational-vibrational non-equilibrium is related to energy transfer behaviour and the translational-vibrational relaxation time. When kinetic energy and chemical energy are converted into internal energy, there is a significant difference in the degree of conversion to translational and vibrational energy. If the translational-vibrational relaxation time is larger than the flow time, such as the relaxation time of the mainstream aftershock wave is 0.25 s for the condition with inflow Mach 2, and the flow time is 3 × 10−5 s, non-equilibrium will occur. Significant differences exist between the flow fields with Mach 2 and 6. A clear boundary layer separation occurs at Mach 6. Combustion occurs at the shear layer, which is in translational-vibrational equilibrium, and there are varying degrees of non-equilibrium in other locations. The dissociation of N2 and production of NO primarily occur on the strut walls and the upper/lower walls of the combustor. The mass fraction of NO is higher than the value at Mach 2. The combustion performance is influenced by the thermochemical non-equilibrium effect. At the condition of Mach 2, it increases the combustion efficiency by 10% near the injector and 0.27% at outlet relatively. Non-equilibrium inhibits the initial upstream combustion while slightly promoting downstream combustion under inflow Mach 6 condition.

This study presents a comprehensive analysis of the frequency response characteristics in a gas generator cycle liquid rocket engine, employing modular decomposition and linearised frequency-domain modeling to simulate dynamic behaviours under forced oscillations. The engine is dissected into key subsystems, including liquid pipelines, turbopump assembly, valves, flow regulation components, thrust chamber, gas generator and pyrotechnic starter, highlighting features such as centrifugal pump pressurisation, staged combustion and cavitation mitigation via venturis. Three oscillation scenarios are examined: supply system responses to thrust chamber pressure disturbances, combustion component responses to fluid disturbances and combustion component responses to pump speed disturbances. Simulations over 0–2000 Hz reveal acoustic-dominated traits in the thrust chamber with oxidiser pathway dominance, low-frequency emphasis in the gas generator driven by fuel disturbances, and heightened instability risks from pump pulsations. Parametric analyses demonstrate that increased pipeline lengths shift resonant frequencies downward, elevated injector pressure drops enhance stability margins by 1.6% with a 20% pressure drop increase, and chamber structural/gas parameter variations erode system stability. These insights, validated against benchmark models, inform strategies for mitigating combustion instability, optimising design parameters, and improving reliability in high-thrust propulsion applications.

As electric vertical takeoff and landing (eVTOL) aircraft technology advances, the focus has shifted towards hybrid electric power systems to overcome battery-specific energy limitations. This study introduces a deterministic hybrid power ratio to develop a series gas turbine hybrid electric propulsion system for a generic quad tiltrotor aircraft. First, the failure modes were categorised into two groups based on the primary power component arrangement, and the risks associated with each other were assessed. Three failure modes were identified in typical eVTOL layouts, i.e. one engine inoperative (OEI), one battery pack inoperative (OBI) and one proprotor inoperative (OPI). In addition, for configurations where a single nacelle contained both the battery and motor, a combined OPI+OBI case was considered, thereby acknowledging interconnected risks and extending the scope to four potential failure modes. The study determined the minimum weight of hybrid power systems using tailored deterministic hybrid power ratios based on five proposed sizing rules. In conclusion, the paper proposes an efficient battery layout for lightweight hybrid power systems and an optimal hybrid power system for the eVTOL aircraft, aligned with current battery technology levels.

The electrohydrodynamic force of a surface dielectric barrier discharge (SDBD) has been well-developed for flow control applications during recent decades. In the present paper, a geometrical modification of the SDBD plasma actuator has been applied to induce a vectorised normal flow at the trailing edge of a NACA0015 aerofoil. The pitot-tube velocity measurements of the normal jet along its propagation direction revealed formation of vortices at the centre of the electrode distance played a role in flow control authority of the jet. The aerodynamic operation of the double-SDBD structure as a virtual flap was assessed versus a single counter-flow jet of a floating structure at pre- and post-stall angles of attack at low Reynolds numbers. It was found that at small angles of attack, the steady counter-flow gives the most effectiveness of lift enhancement in low velocity, whereas in the higher velocity the unsteady one results in more efficacy. The efficiency of both steady and unsteady normal jets increased considerably at high angles such that a lift coefficient improvement of 38% was achieved at $\alpha = 14^\circ $. In the higher velocity, the plasma induced vertical flow acts like a Gurney flap, causing lift increase at high angles by affecting the vortical structures at the trailing edge. Evaluating the obtained results recommended employment of the induced normal flow as a virtual flap at high angles of attack in the unsteady actuation mode.

The penetration strategy of hypersonic vehicles in hostile environments is a critical factor in determining their effectiveness in completing reconnaissance or strike missions. Reinforcement learning (RL), as an end-to-end method, exhibits inherent advantages in addressing complex problems. However, existing research indicates that to enhance the efficiency of RL-based strategies, further advancements are necessary to reduce training costs and improve generalisation capabilities. This paper introduces a RL-based cooperative guidance law for multi-hypersonic vehicles, incorporating the estimated remaining time-of-flight and the absolute value of the bank angle obtained through a predictor-corrector method. The observation space and reward function are specifically designed to simplify the complex decision-making problem into a single-value decision problem, thereby reducing computational complexity and training costs. The proposed guidance law integrates the observation space, reward function and action space within the reinforcement learning framework to control flight trajectories, flight time and penetration of no-fly zones, ensuring compliance with multiple constraints. Model training and simulation tests conducted under multiple constraints demonstrate that the proposed approach reduces the training iterations required for the reinforcement learning agent and improves decision-making efficiency. Furthermore, simulations under different no-fly zone distributions confirm the proposed guidance approach’s high generalisation ability.

This paper presents a study in which modelling and simulation have been used to assess the effect of the aircraft lifts on the air flow over the flight deck of the Queen Elizabeth Class (QEC) aircraft carriers and the subsequent impact on helicopter operations. The aircraft lifts can be either raised or lowered, and they can also have aircraft on them. They can therefore significantly alter the geometry of the starboard side of the ship and, potentially, the air flow over the flight deck. The air flow over the flight deck of the QEC was investigated using experimental and Computational Fluid Dynamics (CFD) techniques. To assess how the air flows for the different lift configurations affected a helicopter landing on the flight deck, piloted flight simulation trials were performed in which a test pilot conducted helicopter deck landings in CFD-simulated Green 60 winds with speeds from 10 kt to 40 kt. Pilot assessment showed the operational wind speed limits, across all spots and lift configurations, were 30 kt or 35 kt and that the different lift configurations produced a 5 kt change in the maximum tolerable wind speeds. While the distribution of the workload experienced by the pilot along the flight path was different for the three lift configurations, it was judged that the difficulty of the overall landing task was not sufficiently affected to require different limiting wind speeds for the different lift configurations.

Hydrogen is a leading candidate for zero-emission propulsion in aviation, particularly when stored and utilised in its liquid form. However, key components such as composite cryogenic pressure vessels remain at low Technology Readiness Levels (TRL), requiring further investigation into their structural performance under realistic operational conditions. The present work aims to provide a validated numerical methodology for simulating the thermomechanical behaviour and the progressive damage evolution of composite cryogenic hydrogen tanks. The finite element framework incorporates ply-level failure criteria and stiffness degradation laws to capture intra-laminar damage mechanisms under combined pressure and temperature loads. The modelling approach is validated against experimental data from coupon-level open-hole tension tests and subcomponent-scale composite pipes burst tests, demonstrating strong correlation in terms of failure onset and progression.

The validated methodology is subsequently applied to a demonstrator, comprising a composite liquid hydrogen tank, subjected to three representative loading scenarios: internal pressure, cryogenic temperature and combined cryogenic-mechanical loading. Results reveal that matrix-dominated damage initiates near the cylinder – dome interfaces of the tank and propagates across the laminate, while fibre failure is not observed in the investigated load cases. This suggests that potential hydrogen leakage is the initial critical failure condition that occurs before any other important structural damage of the tank, highlighting the need for appropriate tank design. The performed study contributes to the understanding of structural integrity of composite cryogenic tanks and offers a computational basis for future design and certification efforts in hydrogen aviation systems.

This article presents the implementation of a new monopulse auto-tracking architecture at the Oran ground station (GS). This architecture is based on a metaheuristic particle swarm optimisation (PSO) algorithm, which measures and adjusts the Q(R) summation associated with the satellite’s main beam direction to ensure optimal synchronisation between the GS and the satellite in terms of antenna pointing. This implementation was validated through practical tests during ALSAT-2B satellite flybys, comprising two distinct scenarios. In the first scenario, the satellite captures new images while simultaneously transmitting data from previously recorded images, thus leading to a misalignment between its antenna and the GS. The second scenario focused solely on data transmission; the satellite being directly aligned with the GS. The results indicate that the pointing error accuracy remains below 0.6 degrees, in accordance with the nominal specifications, thereby enhancing communication performance with a higher received signal level of −55 dBm, which resulted in no loss of images.

The noise levels produced by high-performance supersonic military aircraft engines significantly exceed those of civilian aircraft, highlighting the critical importance of predicting military aircraft noise levels. This paper shows the modeling and assessment of the correlation between sound frequency and sound pressure level (SPL) using particle awarm optimisation (PSO) and the cuckoo search algorithm (CSA) for the high-performance fighter aircraft F-22 Raptor. The developed model aims to predict noise with high precision at various microphone angles from 60° to 150°. As a result of the analysis, the MAPE value for CSA was found to be below 1% for 10 different inlet angles, while the maximum mean absolute percentage error (MAPE) for PSO was 1.7863%. The large dataset ranging from 238 to 762 data points are used and the minimal error values confirm the high accuracy of the model. Additionally, the PSO and CSA algorithms were compared indirectly. The lower error values for CSA, along with its correlation coefficient (R) values closer to 1 indicate that the CSA method gives better results than PSO.

This paper studies a distributed fixed-time dynamic event-triggered formation control framework for a group of hypersonic gliding vehicles (GHGVs) suffering from internal uncertainties and non-affine properties. The main challenge is strong coupling of non-affine nonlinear dynamic with hypervelocity characteristics and multi-source uncertainties make it difficult to design the control protocol. Firstly, by integrating the distributed consensus control strategy, fractional order control theory and dynamic event-triggered mechanism, a framework of fixed-time formation control for GHGVs system is constructed. Secondly, to mitigate the issue of ‘explosion of complexity’ (EI), a fixed-time command filter (FCF) is proposed and a compensative strategy is formulated to tackle the impact of filtering errors. Thirdly, an additional auxiliary differential equation (ADE) is developed to decouple the control input from the status variable. Several radial base function neural networks (RBFNN) are utilised to handle the unknown internal uncertainties. Furthermore, a unique dynamic event-triggered mechanism (DTEM) is introduced for each follower, facilitating seamless transitions between two distinct dynamic threshold strategies. Analysis based on Lyapunov function illustrates that the output tracking error of followers exponentially converges to a small range within a fixed time, and Zeno behaviour is prevented. Finally, several numerical simulations are presented to demonstrate the practicability and meliority of the suggested approach.

This study presents the design and evaluation of an autonomous landing system for rotary-wing unmanned aerial vehicles (UAVs) targeting moving platforms. The proposed system leverages the UAV’s onboard positional data and the moving platform’s position, velocity and orientation information to execute high-precision landings. By incorporating the GPS coordinates provided by the mobile platform, the operational envelope of the landing procedure is significantly extended. A Kalman filter is employed to fuse the platform’s GPS data with the UAV’s inertial orientation measurements, enabling accurate estimation of a dynamic rendezvous point along the platform’s trajectory. This facilitates the generation of an optimised landing trajectory that minimises path length and enhances energy efficiency.

A YOLOv8-based object detection model is integrated into the system to detect the landing pad in real time. The effectiveness of the proposed method is validated through scenario-based simulations designed to evaluate landing performance under variable altitudes, crosswind disturbances and limited visibility due to fog.

Across 30 independent runs, the proposed method reduced total autonomous landing time by 12% (191 ± 2.0 s → 168 ± 1.6 s, p < 0.001), halved the landing phase (22.9 ± 0.7 s → 11.1 ± 0.7 s), shortened the path by ≍152 m (2035 ± 6.8 m → 1883 ± 3.1 m), and lowered battery consumption from 5.0 ± 0.1% to 4.0 ± 0.1%. The system maintained successful landings under variable wind (up to 6 m/s) and fog with a 7 m detection range, achieving sub-meter touchdown accuracy (RMSE ≍ 0.15 m); at a 5 m detection limit, landings failed, indicating a robustness boundary.

Compared to existing literature, the developed system introduces a novel 3D trajectory planning approach involving altitude variations and dynamic target prediction. The framework is modular and compatible with various UAV and ground vehicle platforms, making it suitable for diverse mission profiles in both civilian and defense applications.

Design optimisation of hybrid airships consisting of multi-lobed configurations is being projected as the next paradigm shift in implementing sustainable flight operations within the aeronautical industry. To that end, this paper discusses the effect of varying the relative placement of side-lobes pertaining to a tri-lobed airship hull geometry with respect to the middle-lobe. The paper makes use of a validated OpenFOAM® solver to underscore the aerodynamic impact of shifting the side-lobes in upstream, downstream, upward and downward directions with respect to the middle-lobe while retaining the same volume. These tri-lobed airship hull variants called as skewed tri-lobed hulls, have been comprehensively investigated through the usage of numerical solver (Reynolds-averaged Navier-Stokes) at high Reynolds number flow across various angles of attack. The investigations delineate significant impact of skewed side-lobes on the overall aerodynamics of the tri-lobed airships. Skewing the side-lobes in fore direction leads to drag mitigation at the expense of degraded aerodynamic efficiency owing to lift reduction. Contrarily, aft-skewness amounts to an aerodynamic efficiency enhancement of $ \approx 17{\rm{\% }}$ and improved pitch stability while marginally increasing the pressure drag liability. Aerodynamic efficiency enhancement is attributed to increased lifting force. Skewed upward and downward variants present an overall aerodynamic efficiency reduction. The paper further made use of detailed flow-field visualistion as well as pressure-coefficient distribution plots to underscore the underlying flow-physics related to aforementioned aerodynamic trends. These investigations emphasised the presence of varying three-dimensional relieving effect, intermixing between the three lobes as well as diverse flow separation characteristics downstream of the maximum diameter region leading to the aerodynamic variations thereof. The paper enhances aerodynamic understanding related to tri-lobed geometry that will be crucial in implementing future design changes to the baseline model for improved aerodynamic performance. Amalgamation of these inferences with an optimisation scheme could be implemented in future aerodynamic investigations to optimise tri-lobed geometry for enhanced aerodynamic utility.