This paper presents the concept of a lifting-wing quadcopter unmanned aerial vehicle (UAV), a vertical take-off and landing vehicle (VTOL) with a rear wing, a canard at its front and four propellers. The aerodynamic surfaces are designed so that their mounting angle can be adjusted and fixed before flight, so its performance in transition flight can be studied for a combination of wing and canard mounting angles. A dynamic model using rigid-body equations of motion is presented, which is used to compute the transition flight trajectory from hover to cruise in horizontal flight. The trim conditions were computed for a range of fixed wing and canard mounting angles to study the effects of these variables on transition trajectory parameters such as required power, body pitch angle and propeller rotation speeds as a function of flight speed. Furthermore, a transition flight control algorithm is presented, which has a cascaded PID controller and a reference scheduler to switch between the proper reference states, controls and control allocation matrix. Finally, the transition control algorithm of the conceptual UAV is numerically simulated. Results show that this configuration can perform a fast and smooth transition from hover to cruise flight using the proposed flight control algorithm, substantially reducing required propulsive power in cruise of up to 64%. The application of the control algorithm made notable a transition manoeuver that consists of negatively inclining the aircraft at a negative pitch angle, initially at high intensity, and as the final cruising speed approaches, the inclination is attenuated until the equilibrium pitch angle is reached. Simultaneously with the negative inclination of the pitch angle, there is a slight drop in altitude, which is quickly resumed as the trajectory develops until the final cruising speed. Lastly, this aircraft configuration can be widely used in applications where performance gains in operations currently carried out by multicopters, which cover large distances and need long flight time, would bring great operational advantages.

Current literature offers limited mass estimation methodologies and their application in the conceptual or preliminary design stages of moderate to high aspect ratio wings with electric, hydrogen or distributed propulsions. This study presents the development and application of a quasi-analytical wing mass estimation method to address this limitation. The proposed method is distinguished from the existing mass estimation methods by its expanded realistic load cases, sensitivity to several design parameters, improved accuracy with short computational time and capabilities for future applications. To achieve these features, new geometric models are introduced; 483 load cases including symmetric manoeuvre, rolling, and combined cases are covered following airworthiness requirements; the structural elements are idealised and sized with strength and buckling criteria; existing methods are evaluated and integrated cautiously for secondary structures and non-optimum masses. A computation time of 0.1s is accomplished for one load case. The developed method achieved the highest accuracy with an average error of -2.2% and a standard error of 1.8% for wing mass estimates compared with six existing methods, benchmarked against thirteen wings of different aircraft categories. The effects of engine numbers with dual- to 16-engine setups and the dry wing concepts on the wing mass are investigated. The optimised number of engines and their locations decreased the wing mass of the high aspect ratio wing significantly. In contrast, the dry wing design increased the wing masses of all baseline aircraft. The future applications and improvements of the presented method in novel configurations and multidisciplinary designed optimisation studies are explained.

The aeronautical telecommunication network (ATN) aims to provide reliable end-to-end communications even for those including the air-to-ground segment and in particular for data link applications. The existing ATN, known as ATN/OSI, is based on OSI protocols since its first deployment. The OSI model implementation in ATN communicating entities causes great complexity in network management, particularly in terms of Internet network interoperability. Therefore, since 2010, the International Civil Aviation Organization (ICAO) proposed a migration to ATN over Internet protocol suite (IPS), called ATN/IPS. Thus, this research work focuses on specifying the reliability mechanisms required for air ground data link applications in future ATN/IPS. To achieve this, the transport protocols performance is assessed based on simulations using an ATN model developed considering the ICAO standards. The modeled legacy application enables to generate traffic based on real controller-pilot data link communications (CPDLC) log files from French area control centre (ACC). The air-to-ground subnetworks are characterised using time series delay induced from previously modeled VDL Mode 2 data link analysis. As proof-of-concept, CPDLC messages exchange from aircraft to controller and future applications that transmits heavier files from ground-to-board are simulated. Transport protocols performance are evaluated with respect to the most constraining requirements. The simulation results highlighted the limitations of both connection-oriented transport protocol class 4 (COTP4) and TCP. This enabled to provide a preliminary overview of a new QUIC-like reliable protocol that should meet the heterogeneous requirements of the legacy and the eventual future ATN/IPS applications.

This study presents a model that aims to optimise the sequencing arrival aircraft around the terminal manoeuvring area (TMA). The model considers the transit passenger counts of these aircraft and employs the point merge at Sabiha Gokcen Airport. In this study, aircraft were categorised into two groups, namely ‘High or Low Transit Passenger (HTP/LTP)’. Subsequently, multi-objective models were employed to solve the test problems. Weighted sum scalarisation (WSS), conic scalarization (CS), and epsilon constraint (EC) models were utilised to increase robustness and their results were compared with a single-objective optimisation model. This approach aims to provide decision-makers with a variety of outcomes, thus expanding their options. Simultaneously, efforts are made within the model to allow aircraft with HTP counts to have minimal delays. Additionally, emission calculations were conducted to offer a critical perspective on the environmental implications, and the delay results of the multi-objective optimisation (MOO) models underwent statistical analysis.

Actuator faults in unmanned aerial vehicles (UAVs) can have significant and potentially adverse effects on their safety and performance, highlighting the critical importance of fault diagnosis in UAV design. Ensuring the reliability of these systems in various applications often requires the use of advanced diagnostic algorithms. Artificial intelligence methods, such as deep learning and machine learning techniques, enable fault diagnosis through sample-based learning without the need for prior knowledge of fault mechanisms or physics-based models. However, UAV fault datasets are typically small due to stringent safety standards, which presents challenges for achieving high-performance fault diagnosis. To address this, deep reinforcement learning (DRL) algorithms offer a unique advantage by combining deep learning’s automatic feature extraction with reinforcement learning’s interactive learning approach, improving both learning capabilities and robustness. In this study, we propose and evaluate two DRL-based fault diagnosis models, which demonstrate remarkable accuracy in fault diagnosis, consistently exceeding $99{\rm{\% }}$. Notably, under small sample scenarios, the proposed models significantly outperform traditional classifiers such as decision trees, support vector machines, and multilayer perceptron neural networks. These findings suggest that the integration of DRL enhances fault diagnosis performance, particularly in data-limited environments.

With the development of overall design methodologies for hypersonic vehicles and their propulsion systems, nozzles should expand airflow in a short length and provide sufficient thrust. Therefore, the large expansion ratio single expansion ramp nozzle (LSERN) is widely used. The form of the overexpanded flow field in the nozzle is complex, under the conditions of nozzle start-up, low speed and low nozzle pressure ratio (NPR), thereby negatively influencing the entire propulsion system. Thus, the nozzle flow separation pattern and the key factors affecting the flow separation pattern also deserve considerable attention. In this study, the design of SERN is completed using the cubic curve design method, and the model is numerically simulated for specific operating conditions to study the flow separation patterns and the transition processes of different patterns. Furthermore, the key factors affecting the various flow separation patterns in the nozzle are investigated in detail. Results show that the LSERN in different NPRs appeared in two types of restricted shock separation (RSS) pattern and free shock separation (FSS) pattern, as well as their corresponding flow separation pattern transition processes. The initial expansion angle and the nozzle length affect the range of NPRs maintained by the FSS pattern. The initial expansion angle affects the pattern of flow separation, whereas the nozzle length remarkably influences the critical NPR during transition.

The main purpose of this article is to present the nonlinear unsteady behaviour for jet transport aircraft response to serious atmosphere turbulence in cruise flight and to provide the appropriate mitigation concepts for pilots in the pilot training course of the IATA – Loss of Control In-flight (LOC-I) program. The flight data of a twin-jet and a four-jet transport aircraft encountered serious atmosphere turbulence are the study cases for this article. This study uses flight data mining and fuzzy-logic modeling of artificial intelligence techniques to establish nonlinear unsteady aerodynamic models. Since the rapid change of aerodynamic characteristics in turbulence, so the study uses decoupled longitudinal and lateral-directional motion to identify various eigenvalue motion modes of nonlinear unsteady behaviour through digital 6-DOF flight simulation. It is found that the changes of the main flight variables in the aerodynamic scene and flight environment of the two aircraft are different, but the profiles of five eigenvalue motion modes are actually similar. Those similar eigenvalue motion modes can formulate preventive actions related to the flight handling quality for safe and efficient control by pilots to execute the flight tasks. The one with a large drop height during the ups and downs motion between the two is chosen to construct the movement mechanism of nonlinear unsteady behaviours. The assessments of dynamic stability characteristics of nonlinear unsteady behaviour based on the approaches of oscillatory motion and eigenvalue motion modes related to loss of control will be demonstrated in this article. To develop preventive actions, the situation awareness response to the induced mutation of nonlinear unsteady behaviour on the pilot’s operations will be a further research task in the future.

This paper presents the results of reverse-engineering (RE) strategies, surface roughness and computational fluid dynamics (CFD) modelling for a Wren100 micro gas turbine (MGT). Utilising silicone moulds and resin tooling, precise blade geometry capture was achieved for 3D reconstruction allowing for discrete and parametric geometric models to be created. Using these geometries, CFD simulations employing both Reynolds-averaged Navier–Stokes (RANS) and large eddy simulation (LES) models, alongside experimental wind tunnel cascade tests, were used to evaluate these reverse engineering strategies. The results show that while the parametric model captures overall MGT performance with fewer parameters, the discrete model provides enhanced accuracy, highlighting its suitability for detailed aerodynamic analyses. Contrary to initial expectations, surface roughness exhibited a noticeable impact on performance despite the lower Reynolds numbers (40,000), as demonstrated by the CFD model and wind tunnel experiments. The results indicate that surface roughness can reduce laminar separation bubbles on the blade leading edge, delay the onset of transition, and mitigate secondary flow losses. Overall, this study contributes to knowledge advancement in turbine blade reverse engineering and aerodynamics by detailing the impact of surface roughness on performance.

To meet the development needs of aeroengines for high thrust-to-weight ratios and fuel-air ratios, a high temperature rise triple-swirler main combustor was designed with a total fuel-air ratio of 0.037, utilising advanced technologies including staged combustion, multi-point injection and multi-inclined hole cooling. Fluent software was used to conduct numerical simulations under both takeoff and idle conditions, thereby obtaining the distribution characteristics of the velocity and temperature fields within the combustor, as well as the generation of pollutants. The simulation results indicate that under takeoff conditions, the high temperature rise triple-swirler combustor achieves a total pressure loss coefficient of less than 6% and a combustion efficiency exceeding 99%. Under takeoff conditions, the OTDF and RTDF values are 0.144 and 0.0738, respectively. The mole fraction of NOx emissions is 3,700ppm, while the mole fraction of soot emissions is 2.55×10−5ppm. Under idle conditions, the triple-swirler combustor maintains a total pressure loss coefficient of less than 6% and a combustion efficiency greater than 99.9%. The OTDF and RTDF values are 0.131 and 0.0624, respectively. The mole fractions of CO and UHC emissions are both 0×10−32ppm at the calculation limit of Fluent software.

With the rapid expansion of the aviation industry, an increasing number of Close Spaced Parallel Runway (CSPR) airports are either planning or constructing End Around Taxiways (EAT) to alleviate field operation pressures and enhance safety. Taking Shanghai Hongqiao Airport’s typical CSPR EAT configuration as a case study, this research integrates the airport’s current operational status with the anticipated requirements for future structural renovations and increased flight volumes. Various operational scenarios are established, and simulation research on optimising EAT operations is conducted in advance. The simulation study proceeds as follows: first, an AirTOP simulation model is constructed based on Hongqiao Airport’s actual operational construction. Subsequently, leveraging existing operational scenarios, five simulation scenarios are devised by activating EATs at the departure and approach ends of the eastern zone. The merits and drawbacks of these scenarios are thoroughly analysed. The findings indicate that, with escalating flight volumes, the utilisation of EAT for larger aircraft can curtail their holding duration by nearly 8 min, consequently reducing overall arrival holding duration by 6 min. Departures from gates proximate to T1 experience a 3-min reduction in holding duration through the adoption of EAT at the approach end. Despite an increase in taxi distance due to a higher proportion of aircraft taxiing around, the overall taxi time is diminished. Activating EATs at the departure and approach ends of the eastern zone effectively mitigates the adverse effects of heightened flight volumes on field operational efficiency.

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.

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 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.

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.

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.

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.