In this study, we developed an adaptive gain-scheduling algorithm for hypersonic flight vehicles operating across wide altitude-Mach number envelopes. First, we employed a gap metric-based nominal point selection algorithm to establish a linear parameter-varying (LPV) model more accurate than the traditional Jacobian linearisation method. Active disturbance rejection control (ADRC) was then applied to cope with disturbances and uncertainties, and control gains were scheduled using the Guardian maps (GM) method to adapt to the wide envelope of velocity and altitude. The simulation results demonstrate that under all operating conditions, the proposed algorithm can automatically iterate to obtain a gain-scheduling strategy that meets the flying qualities requirements. Notably, the proposed algorithm exhibited an integral of the time absolute error approximately half of that of the traditional ADRC and significantly lower than that of the GM-LQR method in the ascent phase, demonstrating its excellent control performance and robustness.

A quadrotor was modified by adding wings to the frame to directly compare the flight dynamics characteristics as well as the stability and control derivatives of the quadrotor and its biplane tailsitter variant. The on-axis response of the quadrotor and a biplane tailsitter variant were measured through flight tests, and a frequency domain system identification was used for non-parametric and parametric model identification. Identification of the full vehicle dynamics also demonstrated that identifying the motor torque and back-EMF constants from no-load measurements and the remaining motor parameters from a rotor-motor test stand provided the most accurately identified full vehicle model. The motor dynamics were shown to add a pole to the thrust-based responses (roll, pitch and heave), while the torque-based response (yaw) included a pole and a zero. This approach was then used to identify and compare the quadrotor dynamics, tailsitter dynamics and the total impact of canting the motors. It was found that the presence of the wing added pitch damping to the dynamics and pitch stability became negative. The yaw axis saw an increase in yaw damping derivative, and a reduction in the yaw control derivative to the point where it became difficult to control the aircraft. By introducing cant, both the quadrotor and tailsitter saw large increases in the yaw control derivative. Further, the rotor thrust-based moment generation due to cant resulted in the yaw response zero being canceled by the motor dynamics, resulting in a purely first-order yaw response. Neither the wing nor cant produced any change in the lateral and heave axes.

The energy efficiency of emerging aircraft designs plays a key role, not only in reducing environmental impact, but also in reducing operating costs in the anticipated rise in fuel prices. The European Clean Sky 2 project HLFC-Win is investigating the feasibility of hybrid laminar flow control (HLFC) technology integrated into the outer wing leading edge for a long-haul aircraft. HLFC technology reduces aerodynamic friction drag by means of suction of the boundary layer through a micro-perforated skin to achieve laminarity and thereby improving aircraft performance. However, integrating such a system is not without its drawbacks, as the integration has an impact on the geometry, mass, aerodynamics and engine offtakes that need to be considered. Therefore, the aim of this current work is to assess the HLFC system based on a fair, objective and transparent comparison between the HLFC aircraft and an aircraft of the same technology level without HLFC. The assessment of the HLFC system is twofold, firstly estimating the mission-based performance at the overall aircraft level and secondly performing a lifecycle simulation with three scenarios to determine realistic fuel and cost savings. The mission-based performance assessment indicates a block fuel reduction of over 3 % for the design mission which averages 1.6 to 2.5 % considering a realistic route scenario and expected degradation. The economic assessment suggests a dependency on the scenario chosen, ranging from a 0.7 % increase in total cost (in an unfavourable scenario) to almost a 1 % reduction in total cost (in a favourable scenario), equivalent to $15 million saved per HLFC aircraft over its lifetime. These results support the commercial viability of HLFC technology, which offers significant aerodynamic and fuel efficiency improvements and operating cost savings to the aviation industry. Importantly, no critical barriers were identified for the integration of HLFC technology, further underscoring its potential to improve aircraft performance.

A three-dimensional robust nonlinear cooperative guidance law is proposed to address the challenge of multiple missiles intercepting manoeuvering targets under stringent input constraints and thruster failure. The finite-time convergence theory is used to design a distributed nonlinear sliding mode guidance law, ensuring that the system converges in finite time, with the upper limit of convergence time related to the initial state. A nonlinear sliding surface is adopted to mitigate actuator saturation issues. Then, considering thruster failure, a robust cooperative guidance law is further introduced, ensuring mission completion through the reconstruction of the guidance law. The closed-loop system is proven to be stable using Lyapunov theory, and the influence of hyperparameters on the cooperative guidance law is analysed. Additionally, the results of numerical simulations and hardware-in-the-loop experiments demonstrate the effectiveness and robustness of the proposed algorithm in dealing with stringent input saturation and various disturbances.

In this paper, an intricate autopilot architecture is built for a highly nonlinear aircraft. The mode control of the autopilot is illustrated with state transition diagram. The fundamental concept of this architecture is a model-free design, which means that aerodynamic derivatives and model linearisation are not required. However, only the dynamic control allocation (CA) technique involves the use of derivatives for aerodynamic control. With the use of the Incremental Nonlinear Dynamic Inversion (INDI) in this framework, reduced dependency on models can be achieved. The algorithm addresses system nonlinearities. Our study simulates the approach, and the typical terminal arrival phase of the flight. For lateral navigation, a line-of-sight (LOS) based guidance approach is employed. The principle of the LOS is look ahead-based steering and, mainly the cross-track error is primarily used in the literature. The cross-track error is not the only error for establishing the predefined path. The error between the aircraft heading and the path course is required to describe how fast heading change occurs to re-establish the path again. This proposed modified LOS improves not only the path following ability, but also fault tolerance during aileron or rudder malfunctions. In addition, control surface lock-in-place failures, right/left engine flames-out, and turbulence/wind cases are studied. The objective is a safe approach without reconfiguration. The effectiveness of the proposed control architecture is demonstrated with numerical simulations for the NASA Generic Transport Model (GTM) T-2.

Today, innovative vehicle designs explore the possibility of employing distributed propulsion systems with multiple rotors and propellers. Distributed propulsion systems create complex interactional aerodynamics, that remain largely unexplored and not fully understood. This paper presents a high-fidelity aerodynamic analysis of a tip-mounted propeller combined with over-the-wing propellers. Different configurations were tested using fully resolved simulations with the HMB3 CFD solver. The results indicate that interactional effects in all configurations influence the propeller and wing performance. The proposed configuration, featuring a tip-mounted propeller and over-the-wing propellers, produced 3.5% more thrust compared to the configuration with only a tip-mounted propeller, while also enhancing efficiency. Wing performance was also improved, yielding more lift and less drag, resulting in a higher lift-to-drag ratio. These benefits were due to the thrust and power distributions, and of favourable propeller slipstream/wing interactions. The over-the-wing distributed propulsion system resulted in higher pitching moments, suggesting that moment balancing using different thrust settings and propeller installation positions, or the entire vehicle should be considered to pave the way for practical applications.

Aircraft tyres play a critical role in ensuring the safety of aircraft landings. This paper introduces a novel multi-scale analytical method for evaluating tyre impact performance, explicitly studying the effect of damage defects in the manufacturing and service process on tyre landing dynamic performance. Building on this approach, a numerical simulation of aircraft tyre static and impact load scenarios was conducted, followed by experimental validation. The study systematically compares and analyses the effects of void volume fraction, cord volume fraction and material scale factor on the maximum impact force experienced by aircraft tyre. The variations in maximum impact force arising from changes in tyre structural strength, and deformation can be explained by specific parameters. The findings of this research have significant implications for tyre design and engineering, as well as for enhancing the understanding of the factors that influence tyre performance and safety.

Maintaining high-complexity aircraft requires resilient and data-driven maintenance planning. This article presents the efficient task allocation and packing problem solver (ETTAPS), a novel framework that integrates predictive analytics and optimisation models to generate adaptive maintenance schedules. ETTAPS employs a trial-and-error approach to optimise maintenance intervals, leveraging a branch-and-cut solver combined with first-fit decreasing (FFD) task grouping to minimise costs and enhance aircraft availability. Additionally, a random forest model, retrained using a rolling 24-month data window, continuously refines predictions, leading to progressive cost reductions and improved system reliability over multiple maintenance cycles. Our results demonstrate that ETTAPS significantly reduces maintenance costs and increases aircraft availability by efficiently grouping tasks and incorporating real-world constraints, such as mechanic skill levels, task dependencies and resource limitations. The framework addresses key gaps in MSG-3 and certification analysis, improving task scheduling efficiency and ensuring long-term operational resilience. Furthermore, ETTAPS lays the groundwork for integration with digital twins, real-time anomaly detection and flight planning systems, supporting a more intelligent and proactive approach to aircraft maintenance. This research advances resilience and sustainable aviation maintenance planning by optimising costs, reducing downtime and proactively adapting to operational demands. By aligning with Industry 4.0 and aviation sustainability goals for 2050, ETTAPS contributes to the next generation of intelligent maintenance systems.

Ensuring safety and security in urban air mobility is of utmost significance. As air traffic becomes more concentrated in urban regions, instances of flight conflicts are on the rise. The complex urban morphology, micro-environmental factors and various flight risks significantly impact flight safety. This paper introduces a comprehensive framework to address these challenges. The framework incorporates random 3D city layouts, encompassing city buildings and terrains to establish a corridor graph structure. Leveraging this graph, a multi-view representation learning approach is proposed, which employs graph neural networks, recurrent neural networks (RNN) and contrastive learning to effectively manage tactical conflicts. Through rigorous testing across diverse scenarios, including the incorporation of uncertainties such as wind turbulence, the model’s performance is extensively evaluated. The conclusive results underscore the robustness and efficacy of the proposed approach in ensuring safety and resolving conflicts within the dynamic urban air mobility landscape.

In model-based diagnostics, a simulation model is used to simulate the same operating conditions as the system to be diagnosed to detect and identify anomalies. For this type of analysis, the diagnostic results may be affected by multiple sources of uncertainty. The most common uncertainty to consider is measurement noise. Other sources of uncertainties may originate from the simulation model, instrumentation setup and numerical issues, such as tolerances. While these are often overlooked, they may affect the result to various extent.

In this paper, a multi-point model-based gas path analysis method is proposed and evaluated in the presence of both measurement noise and model uncertainties. The multi-point algorithm addresses the issue of the diagnostic system being underdetermined, having more health parameters than measurements available for diagnostics. It obtains a unique solution through an optimization, where the deviation in health parameter estimation for the operating conditions going into the analysis is minimised. Model uncertainties are introduced in the system by intentionally skewing the characteristics of the rotating components. The objective function is then reconfigured with a, for the gas turbine diagnostic field, novel method taking model uncertainties of the component maps into account. Through this it is possible to reduce the effect of model uncertainties on the diagnostic result. The study shows that through this approach, the uncertainties in diagnostic results are reduced by $3.7{\rm{\% }}$ for the evaluated operating conditions.

Fuel pre-injection in the inlet of a hypersonic engine has been proven to be advantageous in the range of the very high flight Mach numbers. In this paper, a rapid inlet performance analysis model with fuel pre-injection is proposed. The modelling process is divided into two stages. Firstly, the baseline inlet model is provided based on the working principle of the inlet. Then, the newly proposed fuel injection and heat release model is added to the baseline inlet model. Among them, the fuel injection and heat release model is equivalent to increasing the compression angle in the cold state. And in the hot state the effect of the fuel heat release will be considered in addition to the effect of cold state. The research results show as the equivalence ratio increases, the equivalent compression angle also increases, but the two are not in a linear relationship. Based on this pattern of effect, fuel injection can be used to regulate the shock wave position and accurately control the flow rate of the inlet. In addition, by comparing to numerical simulation, it is found that the analysis model can almost reasonably predict the performance of the pre-injection inlet. However, the calculation of drag coefficient has some deviation compared to numerical simulation, which is probably due to the lack of consideration of friction drag and the interaction between the shock wave and boundary layer in the model analysis. Overall, the modelling method proposed in this paper can reflect the effect of fuel injection on inlet performance, which can be used to optimise injection strategy in the future.

Supersonic intakes, under adverse operating conditions, can have unwanted oscillations of shock system internal and external to its duct, known as intake buzz. These buzz instabilities degrade intake performance by causing violent pressure fluctuations, reducing mass flow, decreasing thrust, and leading to combustion instabilities as well. This study examines the onset of this buzz in an axisymmetric intake with various throttles and investigates the effects of dynamic angular deflection of a portion of the cowl on the resulting buzz phenomena. Computations and experiments were conducted at Mach number of 2.0 to obtain the buzz in an axisymmetric intake and investigate its behaviour under various throttle conditions. Dailey-type buzz is observed to be predominant for the present axisymmetric intake, and it has been also quantified that the time period of oscillations for higher throttling ratios is not constant. A technique of dynamically varying the portion of the cowl tip about a pivot point was attempted here to eliminate the intake buzz during onset as well as during a complete buzz cycle. It has been found that the current technique is useful in seizing the buzz shock expulsion from the intake duct, hence restricting unstart and further adversities.

The ground effect phenomenon caused by helicopters in proximity to the ground results in helicopters experiencing a distinctive phenomenon known as brownout in a multiphase environment. However, the substantial computational volume associated with current numerical simulations conducted using the coupled CFD-DEM method restricts the scope of current studies on brownout to individual cases. Consequently, it is currently not feasible to make realistic predictions regarding the impact of rotor design parameters on brownout. In order to make full use of the conclusions of the existing theoretical studies and, at the same time, to save computational resources as much as possible, this paper proposes a novel approach of brownout prediction based on the analysis of the helicopter ground effect flow field eigen quantities in order to gain insight into the nature of the phenomenon of brownout. Firstly, a new approach for predicting helicopter brownout is constructed for the well-developed late-stage ground effect flow field. This is achieved by analysing the rotor flow field characteristics and combining the Greely-Iversen expression in particle dynamics to extract the eigen quantities of each region of the flow field. Secondly, the results of the flow field calculations at different heights are analysed using the aforementioned approach. The effectiveness of the approach is demonstrated by comparing the results with those of CFD-DEM calculations. Ultimately, the results of the numerical simulation of the flow field, when combined with the established prediction approach, allow for the prediction of the brownout phenomenon generated by multiple blade tip shape rotors with different design parameters. Furthermore, a comparative study of the influence of blade tip vortex strength on the development of brownout is conducted, which demonstrates that the rotors of the backswept blade tip have been observed to exert a certain positive effect on the inhibition of brownout, although this influence is limited. In contrast, the rotors of the anhedral blade tip have been seen to transport smaller but larger sand particles with greater efficiency and to re-enter the brownout cycle with greater directness. The rotors of the forward-swept blade tip have been found to cause larger sand particles to participate in the brownout, while simultaneously weakening the transport capacity, which has been resulted in a reduction in the overall degree of brownout.