In fighter pilot training, much of upgrade pilots’ (UPs’) learning takes place during mission debriefs. A debrief provides instructor pilots (IPs) the opportunity to correct situation awareness (SA) upon which the UPs base their tactical decisions. Unless the debrief is conducted with proper depth and breadth, the IPs’ feedback on UPs’ SA and tactical decision-making may be incomplete or false, resulting in poor, or even negative learning. In this study, a new debrief protocol based on the Critical Decision Method (CDM) is introduced. The protocol specifically addresses the SA of UPs. An evaluation was conducted to examine if a short CDM training programme to IPs would enhance their ability to provide performance feedback to UPs regarding their SA and tactical decision-making. The IPs were qualified flying instructors and the UPs were air force cadets completing their air combat training with BAe Hawk jet trainer aircraft. The impact of the training intervention was evaluated using Kirkpatrick’s four-level model. The first three levels of evaluation (Reactions, Learning and Behaviour) focused on the IPs, whereas the fourth level (Results) focused on the UPs. The training intervention had a positive impact on the Reactions, Learning and debrief Behaviour of the IPs. In air combat training missions, the UPs whose debriefs were based on the CDM protocol, had superior SA and overall performance compared to a control group.

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.

Free route airspace (FRA) are increasingly widespread in European countries. A step further, cross-border FRA aims to implement FRA in a seamless,defragmented airspace in Europe. This study aims to bring a different perspective by extending the scope of cross-border FRA applications beyond socioeconomic status (SES) countries. In the study, analyses were made using fast-time simulation outputs for one of the busiest days of Turkish airspace in 2022. The simulation models of four cases were created and compared in such a way that airspaces of Türkiye, South East Europe free route airspace (SEEFRA) countries and Ukraine combined under different border operations above FL285. Flight trajectories were assessed in terms of flight distance, fuel economy and environmental impact within the aforementioned airspaces. The participation of Turkish airspace in the cross-border region and opening of Ukrainian airspace are examined. Concluded, among all, cross-border FRA implemented through all aforementioned regions (Case 4) may constitute the best benefit for flight trajectories and related variables.

Stable separation is a crucial condition that must be met in order for combined aircraft to successfully engage in cooperative flight. In order to achieve the desired fast and controlled separation, this paper proposes a novel design for a torque-driven compliant separation mechanism. By taking into account the compliance characteristics of a sinusoidal acceleration function curve, a mechanical model for the separation mechanism is developed. By utilising the Coulomb friction law, an accurate determination of the aerodynamic load distribution under various conditions is achieved. Subsequently, the relationship between the unlocking moment and the aerodynamic load is derived based on these findings. Through the utilisation of the finite element method, a model of the separation mechanism is generated. To ensure the safety and reliability of the compliant separation mechanism, the mechanical properties of the structural materials are thoroughly analysed under the maximum aerodynamic load. Subsequently, the separation mechanism structure is constructed and subjected to testing in order to showcase the compliance characteristics. In addition, this paper conducts a simulation to analyse the impact of flight speed and angle-of-attack on the separation process. By doing so, the optimal conditions for separation are determined. The methods and findings presented in this study have the potential to contribute valuable insights to the design of combined aircraft.

In this paper, we present a detailed experimental investigation mainly on the vortical flow fields and the associated vortex breakdown phenomena over a non-slender flying wing (sweep angle, {\rm{\Lambda }}${\rm{\Lambda }}$ = 53°). In the process, the aerodynamic coefficients were also determined using a six-component force balance. Surface oil flow visualisation, surface pressure measurements and particle image velocimetry (PIV) measurements, in various crossflow planes and in a longitudinal plane passing through the leading-edge vortex core, were carried out at various Reynolds numbers to understand the flow field over the non-slender flying wing. Aerodynamic characteristics of the flying wing show local peaks and valleys in the pitching moment coefficient. The surface flow visualisation reveals that the nonlinearity of the pitching moment curve is due to the complex nature of vortical flow structures. The flow visualisation also demonstrates the presence of a wave-like surface pattern, and its size is found to reduce with increasing Reynolds numbers. The present PIV measurements confirm that this wave-like surface pattern is associated with vortex breakdown phenomena. These measurements also reveal that the vortex breakdown has not reached the apex of the wing, even at post-stall angle-of-attack. For pre-stall (\alpha $\alpha$ = 20°) flow regimes, it is observed that the location of the vortex breakdown moves downstream as the Reynolds number increases, but this influence is minimised at near-stall (\alpha $\alpha$ = 25°) and post-stall (\alpha $\alpha$ = 30°) flow regimes. Reconstructed velocity field using the first 10 dominant proper orthogonal decomposition (POD) modes reveals that the nature of the vortex breakdown over the flying wing is a spiral-type vortex breakdown.

The aerodynamic performance of a wing model with a row of distributed engines are investigated at the vertical take-off condition. The engines are installed near the trailing edge of the wing. During vertical take-off, the jets exit from the engines and impinge perpendicularly to the ground, providing a thrust for the aircraft. Due to the ground effects, complex vortex structures are induced by the jets. The vortices are categorised into the spanwise vortices and the chordwise vortices. The underwing vortices can lead to low-pressure regions on the lower surface of the wing, resulting in an undesirable downward force. The underwing vortex structures are affected by the ratio of the engine distance to the engine diameter (S/D$S/D$). At a small S/D$S/D$ = 1.10, the flow field is dominated by the spanwise vortices; at a large S/D$S/D$ = 2.78, the flow field is dominated by the chordwise vortices. The range and strength of the spanwise vortices are affected by the vortices interaction. Competition mechanism exists between the range and strength effects, which results in the non-linear variation of the wing lift coefficient with engine spacing. The details of the flow physics underneath the wing and its mechanism on the lift of the wing during take-off are investigated.

The aerodynamic response of a NACA0012 wing section was investigated at a Reynolds number of 100,000 in an open return wind tunnel in the presence of a second wing in tandem. The angle-of-attack of the front wing ranged from −5° to 90° while the rear wing remained at zero incidence. The presence of the downstream wing significantly altered the post-stall behaviour of the upstream wing in the form of a secondary stall characterised by a sudden drop in lift and drag for a specific combination of angle-of-attack and the spacing between the wings. The secondary stall was found to be insensitive to the Reynolds number and the aspect ratio of the downstream wing and did not affect the lift-to-drag ratio. Flow visualisation in the water tunnel indicated that the downstream wing effectively suppressed vortex shedding and lift fluctuations of the upstream wing.

The main purpose of crew resource management (CRM) is to ensure safe flights by preventing possible errors with the effective use of non-technical skills. The aim of the current study is to examine the effects of CRM on flight safety culture (FSC) with the help of the structural equation model with 451 airline pilots. As a result of the analysis, it was determined that there was a significant correlation between CRM and FSC and that CRM has a significant positive effect on FSC. It has been demonstrated that if CRM awareness and skills are used effectively, the perception of FSC will also improve. Furthermore, these findings indicate that there is a need to progress to the corporate CRM phase, i.e., CRM 7.0, to ensure that organisation-wide FSC awareness is established through CRM awareness.

In this work, a three degrees-of-freedom (3-DoF) static quadcopter unmanned aerial vehicle (UAV) test-rig of a pendulum-type configuration is custom-designed, developed, instrumented, and interfaced with a PC. The rig serves as a test bed to develop high-fidelity mathematical models as well as to investigate autopilot designs and real-time closed-loop controllers’ performances. The Simulink Desktop Real-Time software is employed for the quadcopter’s attitude signals acquisition and real-time implementation of closed-loop controllers on a target microcontroller hardware. The mathematical models for pitch, roll, and yaw axes are derived via the first principle and validated with the experimental linear system identification (SI) techniques. Subsequently, employing the multi-parameter root contour technique, the classical proportional integral derivative (PID) controllers are designed and implemented in real-time on the quadcopter UAV test rig. This served as a benchmark controller for comparing it with an integral-based linear quadratic regulator (LQR) controller. Further, to improve the transient response of the LQR controller, a novel robust integral-based LQR controller with a feedforward term (LQR-FF) is implemented, which shows much superior performance than the benchmark and basic LQR controller. This work thus will act as a precursor for a more complex 3-DoF autopilot design of an untethered quadcopter.

With the wide application of quadrotor unmanned aerial vehicles (UAVs), the requirements for their safety and reliability are becoming increasingly stringent. In this paper, based on the feedback of airframe performance health perception information and the predictive function control strategy, the autonomous maintenance of a quadrotor UAV with multi-actuator degradation is realised. Autonomous maintenance architecture is constructed by the predictive maintenance (PdM) idea and the Laguerre function model predictive pontrol (LF-MPC) strategy. Using the two-stage Kalman filter (TSKF) method, based on the established UAV degradation model, the aircraft state and actuator degradation state are predicted simultaneously. For the predictive perception of system health, on the one hand, the system health degree (HD) based on Mahalanobis distance is defined by the degree of airframe state deviation from the expected state, and then the failure threshold of the UAV is obtained. On the other hand, according to the degradation state of each actuator, a comprehensive degradation variable fused with different weight coefficients of multiple actuators degradation is used to obtain the probability density function (PDF) of remaining useful life (RUL) prediction. For the autonomous maintenance of system health, the LF-MPC weight matrixes are adjusted adaptively in real-time based on the HD evaluation, to achieve a compromise balance between UAV performance and control effect, and greatly extend the working time of UAV. Simulation results verified the effectiveness of the proposed method.

Aerodynamic design of a high-efficiency two-stage axial turbine is carried out using a hybrid method through implantation of a two-step design procedure. In the first step, the well-known streamline curvature (SLC) and free vortex (FV) methods are properly combined to establish three-dimensional geometries of the blades at each row and to obtain the flow field properties. The second step is provided to obtain the highest aerodynamic efficiency by optimum clocking of the second stator blades relative to the first ones through executing steady and unsteady computational fluid dynamics (CFD) of three-dimensional viscous flow. Slight discrepancies were observed between gas dynamics results of the SLC and those of CFD. Total pressure and temperature at the turbine outlet, obtained from SLC method, differed from those obtained by 3D-CFD technique by 13.06% and 1.88% respectively. Aerodynamic efficiency of the turbine is obtained about 91.83%, based on 3D-CFD. Time-averaged results showed that under the optimum clocking of the second row stator blades, inlet total pressure and output power of the second rotor increase by 0.23%, and 0.93%, respectively, in comparison to the worst clocking case. These augmentations resulted in increased total to total efficiency of the second stage by 0.444%. Additionally, the total output power of the two stages increased by 0.71% through the optimum clocking. Modeling the unsteady wake flow trajectory within the blades passages confirmed that all of these beneficial effects happen if the upstream wake impinges on the leading edge region of the second stator blades.

The attitude-tracking problem of hypersonic morphing vehicles (HMVs) is investigated in this research. After introducing variable-span wings, the optimal aerodynamic shape is available throughout the entire flight mission. However, the morphing wings cause significant changes in aerodynamic coefficients and mass distribution, challenging the attitude control. Therefore, a complete design procedure for the flight control system is proposed to address the issue. Firstly, the original model and the control-oriented model of HMVs are built. Secondly, in order to eliminate the influence caused by the multisource uncertainties, an adaptive fixed-time disturbance observer combined with fuzzy control theory is established. Thirdly, the fixed-time control method is developed to stabilise hypersonic morphing vehicles based on a multivariable sliding mode manifold. The control input can be obtained directly. Finally, the effectiveness of the proposed method is proved with the help of the Lyapunov theory and simulation results.

For the launch vehicle attitude control problem, traditional methods can seldom accurately identify the fault types, making the control method lack of pertinence, which largely affects the effect of attitude control. This paper proposes an active fault tolerant control strategy, which mainly includes fault diagnosis and fault tolerant control. In the fault diagnosis part, a small deviation attitude dynamics model of the launch vehicle is established, Kalman filters with different structures are designed to detect and isolate faults through residual changes, and the fault quantity of the actuator is further estimated. In the fault tolerant control part, the following control scheme is adopted according to the above diagnostic information: when the sensor fault is detected, the sensor measurement data is reconstructed; when the actuator fault is identified, the control allocation matrix is reconstructed. Simulation results show that the proposed method can effectively diagnose sensor fault and actuator faults, and significantly improve attitude tracking accuracy and control adjustment time.