The aerodynamic performance of an ultra-high aspect ratio strut-braced wing design is assessed for flight at cruise. The sensitivity of a selected airframe design from a recent CleanSky2 project to operating conditions around the design point is quantified using the adaptive-cut high-dimensional model representation (HDMR) method, which allows for the decomposition of the parameter space into smaller subdomains to isolate the parameter interactions and influence on the aerodynamic forces. A comparative analysis with a cantilever wing configuration is performed to identify the role of the strut on the sensitivity of the design. Insight into the transonic performance is gained by characterisation of buffet limits and drag rise. Results show that, for the selected optimised airframe configuration, small changes in freestream parameters can lead to significant reduction in performance due to drag divergence triggered by the shock wave generated at the strut-wing junction and at the fuselage-strut intersection. Cruise conditions can be achieved without buffet onset throughout much of the parameter space. Safety margins associated with buffeting are satisfied, but sensible limits are imposed on the flight envelope for this configuration.

This paper investigates a method for estimating emissions from hybrid-electric aircraft, providing a qualitative assessment of their potential environmental benefits. Limited data availability poses challenges, discussed in this study, for its rigorous evaluation. Proper characterisation of both ${{C}}{{{O}}_{{2}}}$ and non-${{C}}{{{O}}_{{2}}}$ emissions across the operating envelope is crucial. Existing databases and predictive models for turboshaft engines for full-thermal regional aircraft may not be reliable for hybrid-electric configurations, which could operate under uncommon conditions, as low thermal power supply during specific flight phases. This uncertainty highlights the limitations of current emission prediction models. The first of the present study examines non-${{C}}{{{O}}_{{2}}}$ emissions (${{N}}{{{O}}_{{x}}}$, HC, ${{S}}{{{O}}_{{2}}}$, CO) from hybrid-electric aircraft optimised for minimum block fuel, comparing them with those from thermal competitors. The second part evaluates ${{C}}{{{O}}_{{2}}}$ emissions, considering both in-flight and electricity generation contributions. Results indicate that hybrid-electric configurations could reduce non-${{C}}{{{O}}_{{2}}}$ emissions for both the entire mission and the landing-take-off cycle. However, ${{C}}{{{O}}_{{2}}}$ reductions are only significant for short design ranges and could be marginal if electricity generation does not shift to fully renewable sources.

The frequency responses of circulation control and separation control using mini-spoilers for loads attenuation on plunging swept and unswept wings were compared in a water tunnel study. At the pre-stall angle-of-attack, the effectiveness of the spoilers significantly diminishes with increasing reduced frequency of the plunging motion. For the leading-edge spoiler, this happens because the roll-up of the vorticity promotes flow reattachment and reduces the effectiveness of loads attenuation. For the trailing-edge spoiler, the effectiveness of lift attenuation also decreases with increasing reduced frequency, due to the shedding of leading-edge vortices and immersion of the trailing-edge spoiler in the separated flow. The decay of the frequency response for both types of spoilers is similar, implying that it is dictated by the flow separation near the leading edge of the wing in both cases. With increasing sweep angle of the wings, the spoilers’ effectiveness decreases significantly in comparison to the unswept wing. Strong spanwise flow develops for the leading-edge spoiler, which sheds a streamwise vortex, with the same direction of rotation as the wing-tip trailing vortex. This causes partial reattachment of the flow and reduction of the separation area behind the spoiler. With increasing reduced frequency, strong leading-edge vortices dominate the flow over the wing. The leading-edge vortices generate additional vortex lift and also cause the trailing-edge spoiler to be immersed in the massively separated flows. Both factors reduce the effectiveness of the spoilers.

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.

Under the coupling effect of node position deviation, joint clearance and wear factors, the complex landing gear retraction mechanism suffers from low kinematic accuracy, slow retraction performance and shortened reliable life. Addressing these issues, a time-dependent reliability analysis and optimisation design method for the kinematic accuracy of the retraction mechanism is proposed, considering the uncertainty of node position deviation, initial clearance, and dynamic multi-joint wear. Initially, a wear prediction model and a dynamic model of the retraction mechanism considering node position deviation and joint clearance are established to analyse their influence on retraction accuracy and joint wear depth. Subsequent retraction testing under various working conditions is conducted to ascertain the critical failure condition and validate the simulation model. The time-dependent kinematic accuracy reliability model, accounting for the dynamic evolution of wear clearance, is then established to assess reliability variation with retraction cycles. Finally, the reliability optimisation design focusing on hole-axis matching accuracy aims to strike a balance between accuracy cost and reliability, thereby enhancing performance and prolonging operational life.

As avionics systems become increasingly complex, traditional fault prediction methods are no longer sufficient to meet modern demands. This paper introduces four advanced fault prediction methods for avionics components, utilising a multi-step prediction strategy combined with a stacking regressor. By selecting various standard regression models as base regressors, these base regressors are first trained on the original data, and their predictions are subsequently used as input features for training a meta-regressor. Additionally, the Tree-structured Parzen Estimator (TPE) algorithm is employed for hyperparameter optimisation. The experimental results demonstrate that the proposed stacking regression methods exhibit superior accuracy in fault prediction compared to traditional single-model approaches.

This paper presents a comprehensive approach for mitigating noise pollution from unmanned aerial vehicles (UAVs) in urban environment through path planning using reinforcement learning (RL). The study focuses on Turin, Italy, leveraging its diverse urban architecture to develop a comprehensive model. A detailed 3D occupancy grid map, based on OpenStreetMap data, was created to represent buildings’ locations and heights while a population density map was developed to account for demographic variances. The research develops a dynamic noise source model that adjusts noise emission levels based on UAV velocity, ensuring realistic noise impact predictions. Acoustic ray tracing techniques are utilised to simulate noise propagation, accounting for atmospheric absorption and reflections from urban structures, providing a detailed analysis of noise distribution. The core of this work is the application of the deep deterministic policy gradient (DDPG) algorithm within the RL framework. The algorithm is tailored to optimise flight paths by minimising noise impact while balancing other factors like path length and energy efficiency. The RL agent learns to navigate complex urban landscapes, integrating penalties for idling, excessive path length and abrupt manoeuvers to refine its path planning strategy. Simulation results with several maps unseen during training reveal that the RL-based approach effectively reduces noise impact in urban settings, making it a viable solution for better integrating UAVs into urban air mobility (UAM) systems. The methodology is scalable and adaptable, with potential applications in various urban environments globally. This research contributes to the development of sustainable drone operations in UAM context by addressing the critical issue of noise pollution, enhancing public acceptance and regulatory compliance.

This paper quantifies the impacts of the airframe configuration change on the performance differences between a tube-and-wing and a blended wing body aircraft. Both are sized for a 5,000 nmi design range carrying 225 passengers, initially using the same engine. Parametric geometry is created for both concepts based on relevant public information. The tube-and-wing notional geometry is derived from the existing Boeing 767-300ER, whereas JetZero’s concept inspires the blended wing body. These geometries are optimised using computational fluid dynamics and gradient-free approaches. Drag polars for each optimised model, spanning the expected operating envelope, are generated using computational fluid dynamics simulations and multi-fidelity surrogate models. Mission analysis is performed for the blended wing body, a conventional tube-and wing variant with metallic structures, and an advanced tube-and-wing with composite structures. The results show that the blended wing body operates with 15-20% higher lift-over-drag during the cruise, 24% lower fuel burn for the design mission, and 15% reduction in ramp weight relative to the conventional tube-and-wing. These differences drop to 20% for the design mission fuel burn and 10% for the ramp weight relative to the advanced tube-and-wing. When the engines are re-sized and optimised separately for each configuration, the blended wing body demonstrates a 25% improvement in block fuel and 16% reduction in ramp weight relative to the conventional tube-and-wing, which decreases to 21% and 10% relative to the advanced tube-and-wing. In both comparisons, the fuel efficiency advantage of the blended wing body decreases as the mission range is reduced.

Rapid and comprehensive fighter optimisation is an important part of modern combat decision-making. However, due to the numerous influencing factors, it is difficult for decision-makers to consider comprehensively and specify the optimal decision, and it is highly subjective, which leads to different decision conclusions from person to person. Therefore, to solve the above deficiencies in fighter selection, this paper proposes a sequential decision-making framework that comprehensively considers the effectiveness, maintenance, support capability and health status of the fighter aircraft. Based on the multi-dimensional state, it provides comprehensive and credible auxiliary support for commanders. The sequential decision-making framework (called GRA-VIKOR-IFNs) uses the combination of equation and fuzzy multi-criteria decision-making (FMCDM) to evaluate the effectiveness, support capability and health in turn, to complete the step-by-step selection of fighter models, troops and sorties. The evaluation equation is for the effectiveness evaluation and a hybrid method using the extended grey correlation analysis (GRA) and VlseKriterijumska Optimizacija I Kompromisno Resenje (VIKOR) method based on intuitionistic fuzzy numbers (IFNs) is for the support capability and health evaluation. The proposed strategy is in line with the logic and demand of actual combat and training decision-making and takes into account the influence of uncertain factors. Finally, a comparison with some classical methods is carried out, such as the full consistency method (FUCOM), the technique for order of preference by similarity to ideal solution (TOPSIS) and so on. The GRA-VIKOR-IFNs method is consistent with the results of other methods and the result sort resolution is 0.0619 and at least 40% higher than other methods, which can lead the commanders to a more reliable and clear decision.

The impact of the self-sealing band on interior ballistics is investigated during the gun launching, and a high-precision interior ballistics coupling algorithm that takes leakage into account is proposed. This study focuses on a 65 mm short-barrel, equal-caliber balanced cannon, integrating Abaqus finite element software with an interior ballistics calculation programme. It uses a User-defined AMPlication Load (VUAMP) subroutine to achieve real-time coupling calculations of the chamber pressure and self-sealing band deformation, correcting variations in the chamber pressure. Experimental results show that the coupling algorithm offers the higher precision compared to traditional interior ballistics models and can effectively capture the impact of leakage on the interior ballistics performance. Further research reveals that changes in the charge amount and assembly gap significantly affect the sealing performance of the self-sealing band and the leakage of propellant gases, which in turn influence the chamber pressure and projectile velocity. The high-precision coupling algorithm proposed in this paper provides the effective theoretical support for the design of the self-sealing band and the analysis of cannon performance.

This CFD study investigates the aerodynamic characteristics of various internal grid patterns of grid fins without missile bodies, focusing on their influence on aerodynamic performance in subsonic flight regimes. Research examines four distinct grid fin lattices through comprehensive numerical simulations at Mach numbers 0.1, 0.3 and 0.5, for varying high angles of attack ranging between 0o ≤ α ≤ +45o. For validation, comparison between current numerical results and a previous experimental study results was carried out (good agreement within 9.8% error). The numerical results reveal that the diamond cell pattern demonstrates increased lift generation but incurs higher drag, while the square pattern minimises drag and pitching moment, favouring weight-efficient designs. The hexagonal grid fin excels in stall delay and post-stall performance, enhancing manoeuverability and stability at the expense of increased pitching moments. The triangular pattern offers balanced aerodynamic performance with high aerodynamic efficiency at lower speeds, making it versatile for various applications. The analysis offers key insights into grid cell pattern’s aerodynamic effects, aiding grid fin design optimisation for improved aerodynamic efficiency and controllability across various flight conditions and high angles of attack.

The axisymmetric nozzle mechanism is the core part for thrust vectoring of aero engine, which contains complex rigid-flexible coupled multibody system with joints clearance and significantly reduces the efficiency in modeling and calculation, therefore the kinematics and dynamics analysis of axisymmetric vectoring nozzle mechanism based on deep neural network is proposed. The deep neural network model of the axisymmetric vector nozzle is established according to the limited training data from the physical dynamic model and then used to predict the kinematics and dynamics response of the axisymmetric vector nozzle. This study analyses the effects of joint clearance on the kinematics and dynamics of the axisymmetric vector nozzle mechanism by a data-driven model. It is found that the angular acceleration of the expanding blade and the driving force are mostly affected by joint clearance followed by the angle, angular velocity and position of the expanding blade. Larger joint clearance results in more pronounced fluctuations of the dynamic response of the mechanism, which is due to the greater relative velocity and contact force between the bushing and the pin. Since axisymmetric vector nozzles are highly complex nonlinear systems, traditional numerical methods of dynamics are extremely time-consuming. Our work indicates that the data-driven approach greatly reduces the computational cost while maintaining accuracy, and can be used for rapid evaluation and iterative computation of complex multibody dynamics of engine nozzle mechanism.

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.

To investigate the flame acceleration to detonation in 2.0 and 0.5 mm planar glass combustion chambers, the experiments have been conducted utilising ethylene/oxygen mixtures at atmospheric pressure and temperature. The high-speed camera has been used to record the revolution of flame front and pressure inside the combustion chamber. Different equivalence ratios and ignition locations have been considered in the experiments. The results show that the detonation pressure in the 2 mm thick chamber is nearly three times of Chapman-Jouguet pressure, while detonation pressure in the 0.5 mm thick chamber is only 45.7% of the Chapman-Jouguet value at the stoichiometric mixture. This phenomenon is attributed to the larger pressure loss in the thinner chamber during the detonation propagation. As the value of equivalence ratio is 2.2, the detonation cannot be produced in the 2 mm thick chamber, while the detonation can be generated successfully in the 0.5 mm thick chamber. This phenomenon indicates that the deflagration is easily to be accelerated and transformed into the detonation, due to a larger wall friction and reflection. Besides, the micro-obstacle has been added into the combustor can shorten the detonation transition time and reduces the distance of the detonation transition.

A demonstration of a fully onboard method for generating background oriented schlieren (BOS) data on a jet exhaust is presented. Readily available commercial camera equipment is used to capture in-flight imagery of a miniature jet engine exhaust mounted on a custom-built model aircraft. The setup for image acquisition and processing algorithms are described. A new process for registration of images to reduce the degrading effects of vibration and flexure of the airframe are developed and presented along with the underpinning BOS algorithm. Results show that jet flows can be visualised using this technique using a contained system on a single aircraft and demonstrate how a simple technique, such as BOS, can be democratised to such an extent that the cost of conducting in-flight jet measurements can be reduced to the budget of any model aircraft flyer.

Effusion cooling is the state-of-the-art cooling technology for gas turbine hot-gas path components. Typically, effusion cooling holes across the entire combustor liner are aligned with the combustor axis, rendering a nominal zero compound angle between highly directional miniature effusion cooling jets and the main flow direction. The pitch of effusion cooling holes is optimised accordingly. However, the swirling main flow results in a non-zero compound angle and an effectively different pitch from the design. The directional effect of effusion cooling as a result of swirling main flow on the adiabatic film cooling effectiveness (AFE) is a combined effect of a non-zero compound angle and a varied pitch. The current experimental study aims to investigate the isolated effects of compound angle on AFE by excluding the influences of varying pitch. With an improved understanding of the sole effects of non-zero compound angles on AFE, the roles that a varied pitch plays in modifying AFE are further discussed to guide future effusion cooling designs under swirling main flow conditions. Binary pressure sensitive paint (PSP) was used to determine AFE experimentally.

Membrane aerofoils are used for the design of small unmanned air vehicles which have gained interest in the past few years. This paper deals with the nonlinear uncertain aeroelastic analysis of an elastically supported membrane aerofoil. The uncertainties in the aerofoil aerodynamic coefficients are estimated due to five uncertain input parameters, which are the initial tension coefficient, the membrane elastic modulus, the stiffness coefficients of the two supporting springs at the trailing edge and the leading edge, and the fifth parameter is the free stream angle-of-attack. Both static uncertain aeroelasticity and dynamic aeroelasticity for a sinusoidal gust loading are considered. A detailed novel parametric analysis is performed to assess the effect of each parameter. The analysis is carried out using a nonlinear aeroelastic finite element method, which is based on the Theodorsen’s unsteady aerodynamics theory. The polynomial chaos expansion method is used for the uncertainty quantification process and for the sensitivity analysis. Also, the Karhunen-Loéve expansion is used to model the random field of the elastic modulus. The interesting results of the analysis show that the effect of each uncertain input depends on the values of the other parameters and that the initial tension is the key parameter. The type of the probability density functions (or histograms) of the aerodynamic coefficients can vary from a Gaussian distribution to an exponential-like distribution.

Unmanned aerial vehicles (UAVs) with fully electric propulsion systems mainly use lithium-based batteries. However, using fuel cells, hybrid propulsion systems are created to improve the flight time and payload capacity of the UAVs. Energy management and energy-aware path planning are important aspects to be explored in hybrid-propulsion powered UAV configurations. These facilitate optimal power distribution among energy sources and motion planning considering energy consumption, respectively. In the literature, although there are many studies on the energy management of hybrid-powered UAVs and path planning of only battery-powered UAVs, there are research gaps in energy-aware path planning of hybrid-powered UAVs. Additionally, the energy management of hybrid-powered UAVs is usually considered independent of path planning in the literature. This paper thoroughly reviews recent energy-aware path planning for small UAVs to address the listed critical challenges above, providing a new perspective and recommendations for further research. Firstly, this study evaluates the recent status of path planning, hydrogen-based UAVs, and energy management algorithms and identifies some challenges. Later, the applications of hydrogen-powered UAVs are summarised. In addition, hydrogen-based hybrid power system topologies are defined for small UAVs. Then, the path-planning algorithms are classified, and existing studies are discussed. Finally, this paper provides a comprehensive and critical assessment of the status of energy-aware path planning of UAVs, as well as detailed future work recommendations for researchers.

The experimental 7AD rotor blade is assessed for flutter stability in hover to identify the influence of aerodynamic contributions related to blade aerofoil, rotor inflow and wake periodicity on flutter onset. For the aeroelastic analyses, the multibody model is tightly coupled with an unsteady aerodynamic model based on Wagner’s function and related enhancements for the general motion of an aerofoil section considering heave and pitch. The mathematical setup of the approximated Wagner function in state space is extended for axial flow to include unsteady effects related to rotor inflow and wake periodicity. Since the aerodynamic model is based on indicial response functions, a separation of these contributions is possible and allows for the study of their impact on rotor blade flutter. The according flutter results are extracted in terms of frequency and damping behaviour for three test cases that differ in the unsteady aerodynamic model for circulation comprising blade aerofoil, rotor inflow and wake periodicity. As known for articulated rotor blades, also the 7AD blade exhibits a classical bending-torsion coupling. The lowest flutter onset is found for unsteady aerodynamics limited to blade aerofoil, whilst the cases with added rotor inflow and wake periodicity show both the same flutter onset at a 5% larger rotor speed. Here, the influence of rotor inflow plays the major role, since it increases the torsion damping within the critical flutter coupling. Added wake periodicity neither changes frequency nor damping and, hence, does not affect the aeroelastic coupling.

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.