The flow in gas turbines exhibits a highly unsteady, complex and nonuniform manner, which presents two main challenges. Firstly, it introduces instrumentation errors, contributing to uncertainties when calculating one-dimensional performance metrics during rig or engine tests using fixed-location rakes. Secondly, it raises mechanical concerns, including high-cycle fatigue due to blade row interactions in turbomachines and thermal fatigue caused by hot-streaks at the combustor exit. Experimental characterisation of the flow nonuniformity in gas turbines is highly challenging due to the confined space and harsh environment for instrumentation. This paper presents recent efforts to address this issue by resolving the nonuniform flow in gas turbines using spatially under-sampled measurements. The proposed approach utilises discrete probe data and leverages a ‘Fourier-based approximation’ method developed by the author. The technique has undergone preliminary experimental validation involving reconstruction of the total pressure distribution in a multi-stage axial compressor and the total temperature field at the exit of the combustor and high-pressure turbine. Results show that, in the multi-stage axial compressor environment, reconstruction of inter-stage total pressure is achieved using a reduced dataset covering less than 20% of the annulus with reasonably good accuracy. The reconstructed total pressure yields almost identical mean total pressure values at all spanwise locations, with a maximum deviation of less than 0.02%. Additionally, reconstruction of the total temperature distribution at an engine-representative full annulus combustor is achieved using measurements at ten carefully selected circumferential locations. Results show that the reconstructed temperature field successfully captures the primary features associated with combustor exit temperature flow. The reconstructed temperature field yields excellent agreement in the magnitude of radial temperature distribution factor (RTDF) and overall temperature distribution factor (OTDF) predictions to the experiment, with a deviation of less than 0.5% for RTDF and less than 2.5% for OTDF. Lastly, reconstruction of the total temperature distribution at the exit of the GE E3 high-pressure turbine (HPT) is achieved using measurements at eight carefully selected circumferential locations. Results demonstrate remarkable robustness in resolving the temperature profile at the HPT exit with high fidelity, irrespective of the HPT inlet conditions. The initial validation results are promising, demonstrating that the new probe layout scheme and the Fourier-based approximation method enable effective characterisation of flow nonuniformity in gas turbines, thereby providing valuable insights into the complex flow of gas turbine engines.

The vertical motions and buoyancy variations of the two VEGA super-pressure balloons, flown in the middle cloud layer of Venus, are discussed. Using data derived from these 1985 nightside flights, estimates are made of the energy required to operate some alternative balloon platform schemes under consideration for future-proposed Venus-atmosphere in situ science missions. Despite the dissimilarity of these alternative platform schemes, the energy inputs required to operate each scheme on the Venus nightside are found to be similar. Estimates of the associated mass penalties of the associated energy sources are also made. Further investigation of a vertical propulsive assist scheme is recommended.

To enhance the performance of anti-ship missiles cooperative attack, this paper proposes a finite-time trajectory shaping-based cooperative guidance law (TSCGL). Firstly, the cooperative guidance model is established on segmented linearisation of the missile’s heading angle. Then, a trajectory shaping guidance law for a single missile is derived by a weighted optimal energy cost function and Schwarz inequality. On this basis, a finite-time TSCGL is proposed combined with trajectory shaping technology and finite-time theory. The desirable finite-time convergence performance can ensure a simultaneous attack. Through an improved method of time-to-go estimation, it is independent of small-angle assumption and relaxes the launching conditions of the missiles. Additionally, the proposed finite-time TSCGL can achieve better damage performance through energy management. Finally, simulation results demonstrate the effectiveness and superiority of the proposed finite-time TSCGL.

This paper proposed a reinforcement learning-based adaptive guidance method for a class of spiral-diving manoeuver guidance problems of reentry vehicles subject to unknown disturbances. First, the desired proportional navigation guidance law is designed for the vehicle based on the initial conditions, terminal constraints and the curve involute principle. Then, the first-order multivariable nonlinear guidance command tracking model considering unknown disturbances is established. And the controller design problem caused by the coupling of control variables is overcome by introducing the coordinate transformation technique. Moreover, the actor-critic networks and corresponding adaptive weight update laws are designed to cope with unknown disturbances. With the help of Lyapunov direct method, the stability of the system is proved. Subsequently, the range values of the guidance parameters are analysed. Finally, the validity as well as superiority of the proposed method are verified by numerical simulations.

For a class of uncertain systems, a non-overshooting sliding mode control is presented to make them globally exponentially stable and without overshoot. Even when the unknown stochastic disturbance exists, and the time-variant reference trajectory is required, the strict non-overshooting stabilisation is still achieved. The control law design is based on a desired second-order sliding mode (2-sliding mode), which successively includes two bounded-gain subsystems. Non-overshooting stability requires that the system gains depend on the initial values of system variables. In order to obtain the global non-overshooting stability, the first subsystem with non-overshooting reachability compresses the initial values of the second subsystem to a given bounded range. By partitioning these initial values, the bounded system gains are determined to satisfy the robust non-overshooting stability. In order to reject the chattering in the controller output, a tanh-function-based sliding mode is developed for the design of smoothed non-overshooting controller. The proposed method is applied to a UAV trajectory tracking when the disturbances and uncertainties exist. The control laws are designed to implement the non-overshooting stabilisation in position and attitude. Finally, the effectiveness of the proposed method is demonstrated by the flying tests.

With the increase of air transportation, some crossing waypoints (CWPs) are becoming bottlenecks in the operation of air traffic networks. This paper presents a CWP operation optimisation framework based on a two-stage optimisation method. First, we considered the interests of airlines and air traffic controllers and established a flight-level dynamic allocation model for the CWP to minimise the flight-level deviation and the number of flight conflicts. A multi-objective, self-adaptive differential evolution-local search hybrid algorithm was used to solve the model in a parallel computing manner. Subsequently, a flight conflict resolution algorithm based on the Monte-Carlo tree search was designed for flight conflicts that existed after the optimisation. Finally, based on real operation data, four experimental scenarios were constructed, and the air traffic operation simulation system was used for experimental validation. For daily traffic and 1.2 times peak traffic scenarios, the average flight-level deviation reduction rates after optimisation were 53% and 39%, and the successful flight conflict resolution rates reached 89% and 75%, respectively. The experimental results showed that this optimisation framework can effectively balance the number of flight conflicts with the efficiency of flight-level usage and directly improve the capacity of the CWP, which can be used as a reference for air traffic control auxiliary decision support systems.

Water vapour and particles in aero engine exhaust can give rise to condensation trails (contrails) in the wake of aircrafts, and recent studies suggest that persistent contrails and contrail cirrus account for circa 50% of the total aviation-derived radiative forcing (RF). The Schmidt-Appleman criterion is widely used to qualitatively predict the formation of contrails. The criterion indicates that the formation of contrails is affected by both aero engine exhaust and ambient air conditions and can therefore provide the theoretical basis to devise contrail mitigation strategies and further allows quantitative assessment of these strategies. This work focuses on water extraction from the aircraft engine exhaust for contrail mitigation. The fuel water emission index (${\rm{E}}{{\rm{I}}_{{{\rm{H}}_2}{\rm{O}}}}$) is one of the key factors that determines whether persistent contrails form or not. It indicates the amount of water produced for every kg of fuel burnt. Research has indicated that water extraction from the exhaust of the aero engine has been considered for Nitrogen oxides (NOx) reduction, but not for contrail mitigation. Assuming that water extraction is indeed possible, the emphasis of this work will therefore be on understanding how much water is needed to be extracted for contrail mitigation depending on the altitude and the relative humidity (RH), with the aim to carry out a meaningful study on the mitigation of persistent contrails and contrail cirrus to enable a fast and considerable reduction in aviation-derived RF.

The flow characteristics of the plume ejected from a micro-jet engine’s rectangular exhaust nozzle have been studied by conducting experimental and numerical analyses. The radiated infrared signature of a plume ejected from a rectangular exhaust nozzle with a large aspect ratio in a jet propulsion engine is known to be significantly lower than that of a plume ejected from a circular exhaust nozzle. The velocity and temperature distributions, which are the flow characteristics of the jet, were measured to investigate this phenomenon. For this purpose, we installed a circular nozzle and a rectangular exhaust nozzle with an aspect ratio of five to a micro-jet engine. The results showed that the plume spreads wider as it moves away from the nozzle exit and that the velocity rapidly decreases in the case of the rectangular nozzle, contrary to the case of the circular nozzle. Similar tendencies were observed for the temperature distribution and magnitude of the ejected plume. Thus, we concluded that the flow distribution caused by the nozzle shape induces a greater drop in the radiated infrared signature of the plume ejected from the rectangular nozzle than the circular nozzle. Flow analysis was conducted to evaluate the flow in and outside the exhaust nozzle; results similar to those of the experiment were obtained. These results show that the ejecting jet has a greater mixing effect on the air outside when using the rectangular nozzle than the circular nozzle.

The Istanbul metroplex airspace, home to Atatürk (LTBA), Sabiha Gökçen (LTFJ), and Istanbul (LTFM) international airports, is a critical hub for international travel, trade and commerce between Europe and Asia. The high air traffic volume and the proximity of multiple airports make air traffic management (ATM) a significant challenge. To better manage this complex air traffic, it is necessary to conduct detailed analyses of the capacities of these airports and surrounding airspace. In this study, Monte Carlo simulation is used to determine the ultimate and practical capacities of the airport and surrounding airspace and compare them to identify any differences or limitations. The traffic mix, runway occupancy time and traffic distribution at airspace entry points are randomised variables that directly impact airport and airspace capacities and delays. The study aims to determine the current capacities of the runways and routes in the metroplex airspace and project the future capacities with the addition of new facilities. The results demonstrated that the actual bottleneck could be experienced in airspace, rather than runways, which was the focus of the previous literature. Thus, this study will provide valuable insights for stakeholders in the aviation industry to effectively manage air traffic in the metroplex airspace and meet the growing demand.

This paper is concerned with the design and control of a tilt-rotor UAV (TRUAV), with the purpose of simplifying the mechanical structure and transition handling. Previous works on TRUAVs control consider a different controller for each flight mode. Furthermore, two sets of actuators are used: propellers in the VTOL mode, and aerodynamic surfaces in the fixed-wing mode. In this work, a new design that does not contain any control surfaces is proposed. A new control strategy is also presented to accommodate this particularity. Unlike previous works, this strategy uses a single controller to handle both flight phases, making the transition between the two phases no longer an issue. Furthermore, such a characteristic makes the drone capable of following any flight trajectory, which is vital for applications such as the tracking of a ground target. Simulations, conducted on the full nonlinear model of the famous Zagi flying wing drone, showed the effectiveness of the proposed control strategy in tracking a typical trajectory profile with a smooth transition from VTOL to fixed-wing mode.

The Arcanum mission is a proposed L-class mother-daughter spacecraft configuration for the Neptunian system, the mass and volume of which have been maximised to highlight the wide-ranging science the next generation of launch vehicles will enable. The spacecraft is designed to address a long-neglected but high-value region of the outer Solar System, showing that current advances make such a mission more feasible than ever before. This paper adds to a series on Arcanum and specifically provides progress on the study of areas identified as critical weaknesses by the 2013–2022 decadal survey and areas relevant to the recently published Voyage 2050 recommendations to the European Space Agency (ESA).

This paper proposes a fixed-time anti-saturation (FT-AS) control scheme with a simple control loop for the 6-Degree-of-Freedom tracking (6-DOF) control problem of spacecraft with parameter uncertainties, external disturbances and input saturation. Considering the external disturbance and parameter uncertainties, the dynamical model of the tracking error is established. The traditional methods of handling input saturation usually add anti-saturation subsystems in the control system to suppress the impact of input overshoot. However, this paper directly inputs the input overshoot into the tracking error model, thus constructing a modified lumped disturbance term that includes the influence of input overshoot. Then, a novel fixed-time disturbance observer (FT-DO) is designed to estimate and compensate for this modified lumped disturbance. Therefore, there is no need to add the anti-saturation structures in the control loop, significantly reducing the complexity of the system. Finally, an observer-based fixed-time non-singular terminal sliding mode (FT-NTSM) controller is designed to guarantee the fixed-time stability of the whole system. In this way, the convergence time of the proposed scheme does not depend on the system’s initial conditions. Simulation results illustrate that the proposed method keeps the control input within the limit while achieving high-precision tracking control of attitude and position.

In this paper, we investigate the constrained attitude control problem of hypersonic vehicles (HVs). An improved prescribed performance dynamic surface control method is proposed based on an adaptive scaling strategy. Because of the uncertain time-varying disturbances, the controlled state may violate the constraint in the prescribed performance control (PPC) framework. An adaptive scaling strategy is introduced in the PPC method to avoid state violation. The performance function is scaled with respect to the state adaptively. Moreover, a nonlinear disturbance observer is used to compensate the sum of external and other internal disturbances of the system. The proposed method improves the system dynamic performance while ensuring the system robustness. Furthermore, the stability of the closed-loop system is proved by Lyapunov analysis. Finally, numerical simulations are implemented to verify the effectiveness of the PPC method and superiority over other methods.

Machine vision has been extensively researched in the field of unmanned aerial vehicles (UAV) recently. However, the ability of Sense and Avoid (SAA) largely limited by environmental visibility, which brings hazards to flight safety in low illumination or nighttime conditions. In order to solve this critical problem, an approach of image enhancement is proposed in this paper to improve image qualities in low illumination conditions. Considering the complementarity of visible and infrared images, a visible and infrared image fusion method based on convolutional sparse representation (CSR) is a promising solution to improve the SAA ability of UAVs. Firstly, the source image is decomposed into a texture layer and structure layer since infrared images are good at characterising structural information, and visible images have richer texture information. Both the structure and the texture layers are transformed into the sparse convolutional domain through the CSR mechanism, and then CSR coefficient mapping are fused via activity level assessment. Finally, the image is synthesised through the reconstruction results of the fusion texture and structure layers. In the experimental simulation section, a series of visible and infrared registered images including aerial targets are adopted to evaluate the proposed algorithm. Experimental results demonstrates that the proposed method increases image qualities in low illumination conditions effectively and can enhance the object details, which has better performance than traditional methods.

A compressible large eddy simulation (LES) is performed to study a pulsed jet actuator that is used to control a turbulent axisymmetric bluff body wake. The actuator is driven at low-frequency ($f = 200$Hz, $S{t_\theta } = 0.029$) and high amplitude (${C_\mu } = 0.034$). The numerical scheme and a suitable boundary condition for the pulsed jet are validated, showing good agreement with experimental results. A comparison of the velocity boundary condition and the moving boundary condition shows that, in the vicinity of the orifice/slot and in the downstream region, the results from these two methods are identical, while the fluid behaviour inside the cavity shows difference. An analysis of the pulsed jet actuator shows that the phase lag of the cavity pressure is determined by the integration of the diaphragm motion and the pulsed jet. The mean total pressure distribution shows that the total pressure loss is concentrated in the vicinity of the slot. Dynamic mode decomposition (DMD) on the pressure field is used to extract coherent structures which oscillate with the same frequency as that of the diaphragm motion. Some small-scale high-frequency structures are also apparent.

Transient numerical simulations were conducted to investigate the influence of large amplitude and fast impact backpressure on a shock train. The fundamental problem consists of a shock train within a constant-area channel with a Ma=1.61 inflow and a pulse backpressure applied to the outlet. The pressure disturbance in the isolator has an intense forcing-response lag. From the moment of the backpressure peak appearance, it takes 36 times the backpressure duration for the pressure disturbance to reach the upstream end. It moves upstream with time in the form of a normal shock wave. As time progresses, the normal shock degenerates into a $\lambda $ shock and a compression wave behind due to the action of viscous dissipation in the boundary layer. Eventually, a multi-stage shock train is formed. The maximum backpropagation distance is a quadratic function of both the pulse backpressure peak and duration, and the relationship between these variables was determined by fitting. When the integral value of backpressure to time is fixed, reducing the backpressure peak while increasing the duration will reduce the backpressure pulsation at the isolator outlet, which will be more conducive to shortening the maximum backpropagation distance than reducing the duration and increasing the backpressure peak. The values of backpressure peak and duration are obtained from the detonation combustion case, which ensures the authenticity of backpressure characteristics. The relevant research conclusions can provide a reference for the design of the isolator of pulse detonation ramjet.

This paper proposes a methodology to define and quantify the precision uncertainties in aerothermodynamic cycle model comparisons. The total uncertainty depends on biases and random errors commonly found in such comparisons. These biases and random errors are classified and discussed based on observations found in the literature. The biases account for effects such as differences in model inputs, the configurations being simulated, and thermodynamic packages. Random errors consider the effects on the physics modeling and numerical methods used in cycle models. The methodology is applied to a comparison of two cycle models, designated as the model subject to comparison and reference model, respectively. The former is the so-called Aerothermodynamic Generic Cycle Model developed in-house at the Laboratory of Applied Research in Active Control, Avionics and AeroServoElasticity (LARCASE); the latter is an equivalent model programmed in the Numerical Propulsion System Simulation (NPSS). The proposed methodology is intended to quantify the bias and random errors effects on different cycle parameters of interest, such as thrust, specific fuel consumption, among others. Each bias and random errors are determined by deliberately preventing the effects from other biases and random errors. The methodology presented in this paper can be extended to other cycle model comparisons. Moreover, the uncertainty figures derived in this work are recommended to be used in other model comparisons when no better reference is available.

The ground delay program (GDP) is a commonly used tool in air traffic management. Developing a departure flight delay prediction model based on GDP can aid airlines and control authorities in better flight planning and adjusting air traffic control strategies. A model that combines the improved sparrow search algorithm (ISSA) and Multilayer Perceptron (MLP) has been proposed to minimise prediction errors. The ISSA uses tent chaotic mapping, dynamic adaptive weights, and Levy flight strategy to enhance the algorithm’s accuracy for the sparrow search algorithm (SSA). The MLP model’s hyperparameters are optimised using the ISSA to improve the model’s prediction accuracy and generalisation performance. Experiments were performed using actual GDP-generated departure flight delay data and compared with other machine learning techniques and optimisation algorithms. The results of the experiments show that the mean absolute error (MAE) and root mean square error (RMSE) of the ISSA-MLP model are 16.8 and 24.2, respectively. These values are 5.61%, 6.3% and 1.8% higher in MAE and 4.4%, 5.1% and 2.5% higher in RMSE compared to SSA, particle swarm optimisation (PSO) and grey wolf optimisation (GWO). The ISSA-MLP model has been verified to have good predictive and practical value.

Most simple ailerons produce adverse yaw. However, with proper aileron placement and wing twist, an aileron can produce proverse or neutral yaw, eliminating the need for aileron-rudder mixing, differential aileron deflection or Frise ailerons. The relationship between wing planform, aileron placement and lift distribution is studied here for a special class of optimal lift distributions that minimise induced drag for a variety of design constraints. It is shown that a wing employing the elliptic lift distribution will always produce adverse yaw, independent of aileron design or operating condition. However, for wings employing other optimal lift distributions, the ailerons can be placed to produce proverse or neutral yaw. A numerical lifting-line algorithm is used to explore the impact of aileron design on a wide range of wing planforms and lift distributions. Results can be used in the early stages of design to correctly place ailerons with respect to desired roll-yaw coupling.

To enhance the impact resistance capacity and ensure the floatability of aircraft after ditching, the slamming response of three types of aircraft sub-floor structures are investigated including the flat, cylindrical and ellipsoidal under floor. A coupled Finite Element-Smooth Particle Hydrodynamic (FE-SPH) method is employed with focus on non-linear structural collapse in fluid-structure interaction. The material is defined by bilinear elastic plastic law, and the strain rate effect is taken into account. Further, comparison and analyses are performed in terms of acceleration, local pressure and strains at different speeds. Results show that conventional flat sub floor structures perform poorly during ditching due to excessive peak acceleration and pressure. By contrast, the peak acceleration of ellipsoidal under floor is lower at all measured speeds and the pressure on the sides is reduced. Moreover, the ellipsoidal sub-floor with bi-directional curvature generates smaller plastic strain and deflection of skin, demonstrating better mechanical properties in water impact scenarios.