Flight Data Monitoring (FDM) programmes have become a key part of every major airline’s safety management system. They are primarily based on learning from unwanted deviations in flight parameters encountered during normal flight operations. Owing to its unique nature, anomaly detection of FDM presents distinct problem complexities from the majority of analytical and learning tasks. This methodology, while useful, concentrates only on a small part of the operation, leaving most of the data unprocessed, and does not allow for analysing events that had the potential to go wrong but were recovered in time by the crews. This research focused on analysing an FDM dataset of 1332 approaches between January 2018 and July 2022 at Tenerife South Airport (Spain), where there is a known phenomenon of increasing headwinds during the final approach. The flights were clustered using self-organising maps (SOM) by patterns of increasing headwinds, and the clusters were assessed in terms of clustering performance. The clusters were well differentiated. A further comparison between the results from the airline showed that 88 flights were affected by wind shifts, while 27 flights were picked up by the airline. The results demonstrate that SOMs are a meaningful tool for clustering flight data and can complement the current FDM analysis methodology. Combining both methodologies could shift FDM data analysis to look beyond exceedances into what went well, thus shifting the FDM paradigm towards a more safety-II-based method.

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.

During the automatic docking assembly of aircraft wing-fuselage, using monocular camera or dual-camera to monitor the docking stage of the fork-ear will result in an incomplete identification of the fork-ear pose-position and an inaccurate description of the deviation in the intersection holes’ position coordinates. To address this, a quality inspection and error correction method is proposed for the fork-ear docking assembly based on multi-camera stereo vision. Initially, a multi-camera stereo vision detection system is established to inspect the quality of fork-ear docking assembly. Subsequently, a spatial position solution mathematical model of the fork-ear feature points is developed, and a spatial pose determination mathematical model of fork-ear is established by utilised the elliptical cone. Finally, an enhanced artificial fish swarm particle filter algorithm is proposed to track and estimate the coordinate of the fork-ear feature points. An adaptive weighted fusion algorithm is employed to fuse the detection data from the multi-camera and the laser tracker, and a wing pose-position fine-tuning error correction model is constructed. Experimental results demonstrate that the method enhances the effect of the assembly quality inspection and effectively improves the wing-fuselage docking assembly accuracy of the fork-ear type aircraft.

This paper presents a comprehensive evaluation of reduced-order models (ROMs) for the determination of pressure coefficient distributions on supersonic and hypersonic bodies. The study investigates the limitations, aerodynamic precision and computational performance associated with various methodologies, ranging from simplistic Newtonian theory-based approaches to more advanced first and second-order shock-expansion theories. Validation is performed by comparing computed results with experimental and computational data for pressure distributions, drag and lift coefficients and centres of pressure for fundamental geometries and authentic vehicle design over a wide range of freestream conditions. The study also includes a comprehensive computational complexity analysis, demonstrating the superiority of finite-element ROM approaches over traditional finite-volume computational fluid dynamics (CFD) simulations. The primary objective of this paper is to scrutinise the extension of these methodological classes to the low supersonic regime. Hence, thermo-chemical reactions within the flow are disregarded, and the ideal gas law is adopted. A value of $\gamma = 1.4$ is chosen for consistency and comparability across the analyses. The proposed ROMs show remarkable potential for reducing high-speed simulation execution times by four orders of magnitude, maintaining accuracy within 20 per cent and as low as 1 per cent. The study unveils three key findings: first, the accuracy degradation of Newtonian-based theories for inclined elements, particularly around 45 degrees, and their reduced dependency on Mach number at large inclination. Secondly, the study presents novel insights into the impact of shock-wave-Mach-wave interactions on pressure distribution calculations, emphasising the Mach number as a crucial metric governing recompression effects. Lastly, the study demonstrates the exceptional accuracy of DeJarnette’s method, providing ${C_P}$ results within 2 per cent for a wide range of conditions, offering an attractive alternative to the Taylor-Maccoll equation.

This study investigates the influence of release timing on the trajectory of internal store separation through numerical solutions of continuity, momentum, energy equations and six degrees of freedom equations in a coupled manner. The internal store separation process in advanced fighter aircraft is analysed using computational fluid dynamics (CFD) and six degrees of freedom equations of motion. Initially, the equations of motion are validated by reenacting the Eglin Air Force Base study, an external store separation example with documented experimental results. Subsequently, validation is extended to the M219 cavity problem. In the internal store separation analysis, a cavity with an L/D ratio of 5, a freestream velocity of 0.85 Mach, and a generic store are utilised. Detached eddy simulation (DES) is applied using both static and dynamic mesh techniques in all numerical solutions. The generic store, positioned within a clean cavity with a 90-degree flap angle, was released at two distinct times, corresponding to the points of maximum and minimum gravitational forces. Interestingly, the results show that releasing the store when the normal force acting on it is at its maximum does not necessarily provide an optimal separation. Specifically, when the force coefficient was at its maximum (0.14), the store collided with the cavity door flap after 0.171465 seconds. In contrast, when the force coefficient was at its minimum (-0.04), the store contacted the cavity door after 0.170295 seconds at the same location. Despite the differences in force magnitudes, the trajectories were nearly identical, suggesting that the timing of the release may not have a significant effect on preventing collision. This further emphasises the need for flow control methods to ensure safe and effective store separation.

The threat of GNSS interference poses a great danger to many critical infrastructure systems including air navigation. With a focus on mitigating this threat, this paper proposes a methodology for detecting GNSS interference. The methodology utilises the quality indicator NACp transmitted in ADS-B messages and GPS almanac data for interference detection. The NACp indicator enables estimation of the position error derived from GPS, which is compared with the HDOP value of the GPS satellite constellation. Based on this comparison, the developed detection algorithm determines whether the aircraft is affected by jamming. The detection methodology is evaluated on datasets obtained during deliberate experiments with GPS jamming. The proposed methodology provides a way to detect GNSS interference, facilitating mitigation of its impact on air traffic operation.

The presented research investigates the impact of interference on the performance of aircraft Global Navigation Satellite System (GNSS) receivers with a specific focus on the behaviour of Automatic Dependent Surveillance-Broadcast (ADS-B) position quality indicators. Several experiments were performed with different aircraft types, such as Airbus, Boeing, Beechcraft King Air B350 or Tecnam, and using various intensities of GNSS jamming. The behaviour of various quality indicators, such as the Navigation Integrity Category, Navigation Accuracy Category, Source Integrity Level and System Design Assurance transmitted in different types of ADS-B messages, is analysed. We investigate not only situations where the quality indicators drop to zero, but also the complete evolution of the changes in the indicators as a function of the increasing power of the jamming signal. Based on the analysis of changes in the ADS-B quality indicators, the estimation of the most likely interference signal power required to discontinue the tracking of an already acquired GPS L1 Coarse/Acquisition signal is made. Additionally, the interference signal power to prevent re-acquisition is also estimated. The findings improve the understanding of interference effects and can support the development of robust interference mitigation techniques in aviation applications.

A terrain and path following control scheme is designed for ground detection mission of a fixed-wing unmanned aerial vehicle (UAV) considering the attitude constraint. The attitude of the UAV should be maintained for efficient exploration, leading to the degradation of mission performance. The proposed controller makes the attitude of the UAV remain in a desired range, which alleviates the mission performance degradation. The proposed algorithm consists of the guidance law and the nonlinear flight path controller. The guidance law is designed by combining a terrain-following altitude controller and a horizontal path following controller based on the Lyapunov control scheme. The generated command by the guidance law is used as a reference input to be followed in the flight path controller. The flight path controller is designed considering the attitude constraint. Especially, the roll and pitch angles of the UAV are considered as attitude constraints so that the angles remain within the desired range. To design a flight path controller satisfying the attitude constraint, the control system is decomposed into three feedback loops. State-feedback controllers are designed using the sliding mode control scheme for flight path control in the outermost loop as well as for angular rate control in the inner loop. In the second-outer loop, a quadratic programming (QP)-based controller is designed to control the sideslip angle while satisfying the attitude constraint. The control Lyapunov function is adopted to determine the QP constraint for the sideslip angle control, and the control barrier function is used to obtain the QP constraint for the attitude constraint. Numerical simulation is performed to demonstrate the effectiveness of the proposed algorithm.

In this paper the performance of an internal strut-based thrust vector control (TVC) system has been studied at different expansion conditions of propulsion nozzle. The TVC system uses a cylindrical strut inserted through the diverging wall of a supersonic nozzle. This TVC system can be construed as an alternative to secondary injection TVC method. The nozzle had an expansion ratio of 1.545 and nozzle pressure ratio (NPR) of 6.61 for optimum expansion. Numerical simulations were performed at over-expansion (NPR = 3.94) and under-expansion (NPR = 7.89) conditions for four strut locations (xs) and five strut heights (hs). The strut location from the nozzle throat corresponded to 33%, 50%, 66.7% and 80% of the diverging length (Ld) of the nozzle. The schlieren images of the nozzle exhaust and nozzle internal wall pressure distribution from experiments were compared with the results from numerical simulation and the agreement was quite good. Computational results show that introduction of the strut caused a maximum total pressure loss of 1.5% at its maximum height. The calculations also show that $ \pm $4${}^{\circ}$ thrust deflection angle could be achieved using combinations of strut location and strut height over a range of nozzle operational conditions. Thrust vectoring performance of strut insertion TVC was evaluated using a parameter called vectoring performance index (VPI) defined as thrust deflection angle per unit percentage of pressure loss. The maximum VPI was observed when xs=0.5Ld at ${\bar{h_s}} = 0.429$ in both over-expansion and under-expansion conditions. The study reveals that an internal strut based TVC has a good future potential to be developed as an alternate TVC system obviating the requirement of carrying a fluid tank for a system like secondary injection TVC.

The cavities over the re-entry vehicle alter the aerothermodynamic properties, leading to enhanced thermal protection as well as effective aerothermodynamic performance. This paper investigates the estimation of aerothermodynamic properties over a re-entry vehicle with different types of cavities on the frontal face of the vehicle. The direct simulation Monte Carlo (DSMC) simulation of hypersonic flow over the Crew module Atmospheric Re-entry Experiment (CARE) capsule was simulated with the re-entry velocity of 7,422 m/s and the freestream temperature of 225 K at an altitude of 110 km. A transient flow Knudsen number of 0.1 and air consists of 78.09% of ${N_2}$ and 21.91% of ${O_2}$ are used in the simulations. Two types of cavities, namely trapezoidal and the semi-circular cavity on the frontal face of the re-entry vehicle with different length to depth ratios, are analysed. The simulation results show that the recirculation regions are formed at the base of the cavity in the case of a cavity with sharp corners, whereas in the case of a cavity with rounded corners, the recirculation formed at the lip of the cavity for both trapezoidal and the semi-circular cavities. Increasing the length and depth of the cavity leads to smaller decrement in the drag when compared to the capsule without cavity for both trapezoidal and the semi-circular cavities. The heat flux is low for a cavity with the small L/D ratio (L/D = 0.5) for both fixed length and depth for trapezoidal-type cavity, whereas for large L/D ratio (L/D = 1.5) increasing the length of the cavity increases the overall heat flux.

This article aims to introduce a novel synthesis optimisation framework based on a performance prediction model for a space propulsion system. The structure of this research is divided into five steps for optimisation. In the first step, the optimisation problem is defined, and the objective functions, design parameters and constraints of the problem are determined. This system consists of five sub-systems including (gas tank, liquid fuel tank, injector, catalytic bed and nozzle). In this optimisation, the objective functions include increasing the specific impulse and reducing the overall mass of the system. Also, the design parameters and limitations of each subsystem are mentioned in the text. The optimisation space is extracted in the second step using the design of experiments (DoE) tool and the Latin hypercube sampling (LHS) method. In the third stage, using the design points obtained from the DoE, the metamodel of the mass of each of the sub-systems and the metamodel of the specific impulse of the whole system are produced by the Kriging method. In the fourth step, the metamodels produced are optimised using the SHERPA algorithm based on the objective functions of the problem. In the fifth step, the results are derived by comparing two objective functions (minimum mass and maximum specific impulse for the entire system) using the Pareto front diagram. Finally, by comparing the optimal results with an existing thruster sample, it shows that the specific impulse has increased by 6% and the total mass of the system has decreased by 15.8%.

Propulsive fuselage aircraft complement the two under-wing turbofans of current aircraft with an embedded propulsion system within the airframe to ingest the energy-rich fuselage boundary layer. The key design features of this embedding are examined and related to an aero-propulsive performance assessment undertaken in the absolute reference frame which is believed to best evaluate these effects with intuitive physics-based interpretations. First, this study completes previous investigations on the potential for energy recovery for different fuselage slenderness ratios to characterise the aerodynamics sensitivity to morphed fuselage-tail design changes and potential performance before integrating fully circumferential propulsors. Its installation design space is then explored with macro design parameters (position, size and operating conditions) where an optimum suggests up to 11% fuel savings during cruise and up to 16% when introducing compact nacelles and re-scaling of the under-wing turbofans. Overall, this work provides valuable insights for designers and aerodynamicists on the potential performance of their concepts to meet the environmental targets of future aircraft.

The unmanned aerial vehicle (UAV) system for composite vertical take-off and landing (VTOL) is a complex, highly coupled, and nonlinear system which is sensitive to external disturbances and model uncertainties. The composite VTOL UAV system consists of a multi-rotor section and a fixed-wing section. To improve observation accuracy, the compensation function observer (CFO) uses a new structure that includes velocity information. The CFO is utilised to estimate the uncertainty and the external disturbances of the system model, which performs superior estimation accuracy compared to the extended state observer (ESO). In the modeling process of the VTOL UAV, the aerodynamic moment is calculated by means of the cross-product operation of force and force arm, which solves the problem of over-reliance on aerodynamic parameters in the traditional modeling approach. The controlled object is refined by CFO, and model compensation control (MCC) is used to realise the velocity and attitude control of the composite VTOL. The numerical simulation of MATLAB/Simulink and hardware-in-loop simulation (HIL) of Rflysim were implemented, and which were used to compare the MCC, active disturbance rejection control (ADRC), and proportion integration differentiation (PID). The simulation results confirm the superiority of MCC in controlling composite VTOL UAVs in terms of anti-disturbance and tracking speed.

Cost efficiency is a critical factor in the competitive aviation sector. These efficiency factors force airline operators to develop new approaches in their organizations. Predictive maintenance helps to build scheduling maintenance programs for airline operators or MROs. Scheduled maintenance programs benefit cost efficiency in the aviation sector. Predictive maintenance methods predict the failure time of any equipment. Predictions can be made by analyzing the sensor values from equipment.

In this paper, we predicted the remaining useful life (RUL) of turbofan engines using machine learning models and a similarity-based approach. Sensor datasets from the Prognostics Data Repository of NASA, called CMAPPS, were utilized. Using the FD0002 sub-dataset, a health index (HI) was created, and models were trained. Once the models were trained, train and test HIs were estimated. The predicted test HI was matched with the predicted train HI based on a similarity-based approach, and then a RUL prediction was made. The results obtained were compared with the actual results to calculate the accuracy, and the algorithm that resulted in the maximum accuracy was identified.

We selected six machine learning algorithms and also created an ensemble model by averaging the predictions of six machine learning algorithms for comparing prediction accuracy. The different algorithms were compared to obtain the prediction model with the closest prediction of remaining useful lifecycle in terms of the number of life cycles. This experiment showed us the effect of the similarity-based approach on the basic version of machine learning models for RUL prediction.

Thrust changes near walls and the ground plane are influenced by the rotor’s position in indoor flight environments. This study evaluates variations in rotor thrust near a corner, which includes one wall and the ground plane, as well as a vertex, which involves two walls and the ground plane; these phenomena are referred to as the corner and vertex effects, respectively. Additionally, the rotor wake in the vertex effect was visualised using the laser sheet method, and wake velocity was measured with a hot-wire probe. The thrust change in the corner effect on the ground side was minimal, primarily depending on the ground effect. In the vertex effect, thrust decreased to 93% of the thrust outside the vertex effect when the rotor height above ground was 2.5 times the rotor radius, and the rotor was distanced from the two walls by 1.5 times the rotor radius. Flow visualisation and hot-wire velocimetry results suggest that the thrust decrease was caused by the flow recirculation structure between the fountain flow developed along the vertical corner and rotor inflow. The thrust decreases under conditions where the circulation structure appears, and fountain flow velocity accelerates the recirculation. These findings aid in planning the flight path of small multirotors in indoor flight conditions by providing guidance on distances that do not alter rotor thrust near corners and vertices.

Station-keeping control is a critical technology for stratospheric aerostats. For those aerostats that utilise wind field environments to achieve trajectory control, the station-keeping capability of a single aerostat is inherently limited. This limitation can lead to instances of the aerostat flying outside the designated task area, thereby diminishing the effectiveness of station-keeping control. To ensure continuous monitoring of the restricted area for long endurance, dynamic adjustments and cooperative coverage among multiple aerostats are necessary. This paper introduces an optimal coverage algorithm based on Voronoi diagrams and presents a formation control method for stratospheric aerostats that employs the virtual force method and the ${A^{\rm{*}}}$ algorithm, respectively. In a real wind field environment, ten aerostats are deployed to optimally cover the restricted area. Simulation results indicate that the coverage rate of the stratospheric aerostats within the restricted area can exceed 70%, while the network connectivity rate among the aerostats can reach 80% following guidance control during return flights. Furthermore, the stratospheric aerostats that flying out of the restricted area can return through path planning and optimal coverage algorithm, and the networking connectivity rate between aerostats is higher than that using the virtual force method.

This paper presents an improved signal-processing method based on the Hilbert-Huang transform (HHT), which is applied to the fault feature extraction of the aerospace generator rotating rectifier (AGRR). Initially, the excitation current of the alternating-current (AC) exciter is utilised as measurable information for data collection. Subsequently, the HHT is processed with variational mode decomposition (VMD), followed by the improvement of the variational Hilbert-Huang transform (VHHT) using particle swarm optimisation (PSO) to determine the modal decomposition number and the secondary penalty factor. Finally, the proposed PSO-VHHT method is compared with several other signal processing-based feature extraction methods through both simulated and practical experiment data, and an analysis of the diagnostic performance of these methods is also conducted.

For binary plug nozzle, the plug cone is exposed to high-temperature mainstream flow, making it one of the nozzle’s high-temperature components. This paper uses the Realizable k-ε turbulence model and the reverse Monte Carlo method to numerically investigate the aerodynamic and infrared radiation characteristics of the plug nozzle. Various slot cooling configurations were adopted to study the nozzle’s infrared radiation in detail. Results indicate that compared to the baseline nozzle, the plug nozzle’s performance is slightly reduced due to the decrease in effective area of flow over the plug cone. Introducing slot cooling at the rear edge provides significant infrared suppression benefits at low detection angles and notably reduces infrared radiation discrepancy with baseline nozzle at high detection angles. The cooling air from slots causes the nozzle jet to exhibit a ‘thermal layered’ feature. With the same total coolant mass flow, the ‘leading edge + trailing edge’ cooling configuration can lower the area-averaged wall temperature of the plug cone by 5.5% – 12.3%. However, its infrared radiation intensity at each detection angle on the pitch detection plane is higher than that of the ‘trailing edge’ configuration. The significance of leading-edge cooling is focused more on thermal protection for the plug. Thus, it is essential to balance coolant mass flow distribution between infrared radiation suppression and thermal protection.

Typically, the fuselage of a modern military aircraft is designed in such a way that the propulsion system is integrated into it. The main reasons are reduction of installation space and minimisation of radar signature. Those requirements can be achieved by using highly bent engine intakes, which are occluding a direct line of sight to the compressor system. Depending on their design, secondary flows and flow separation can be expected due to the strong curvature of the intake system. In this study, a serpentine intake in front of the Larzac 04 test engine is investigated experimentally and its performance compared with and without flow stabilising measures. In detail, a configuration with vortex generators was compared experimentally with a configuration of active flow control by injected air. In order to analyse and compare the efficiency of both systems, the dimensionless total pressure coefficient, the distortion coefficient (DC60) and detailed surface pressure distributions as well as the aerodynamic interface plane are evaluated. In addition, different throttle lines were recorded for surge line evaluation of the low pressure compressor of the Larzac engine and compared for each flow-stabilising measure investigated. It was found that the application of injecting air showed a larger improvement in surge margin and reduction in distortion coefficients compared to the passive flow control.

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.