To study the performance of the main windshield of a commercial aircraft that has been verified to be airworthy by bird-strike tests against unmanned aerial vehicle (UAV) impact at high-speed, a typical light UAV with various possible flight postures and the main windshield of a commercial aircraft are considered. The transient impact responses at critical moments, energy change and contact force of a multi-layer windshield impacted by a UAV with different postures are investigated using a simulation method based on the models verified by the high-speed impact test between the whole UAV and the full-size nose. This study shows that the flight posture of the UAV has a significant effect on the damage to the windshield. When the abdomen of a typical light UAV maintains a posture parallel to the plane of the windshield, the high-speed impact would cause catastrophic damage to the windshield and no longer be airworthy. Simultaneously, the damage to the aircraft windshield caused by UAV collision is far more serious than that caused by bird strikes under similar collision conditions. The mass-concentrated components of the UAV and their high-hardness characteristics are the main factors of affecting multi-layer glass of windshield damage. The degree of damage to the windshield is positively related to the absorbed energy rather than the impact contact force. In this study, the impact simulation results between the windshield and UAV with different flight postures are verified qualitatively by testing, which provides a rational understanding and technical pre-research support for emerging and increasingly frequent potential safety hazards in air transport practice.

This paper investigates the issue of tracking control for a free-floating space manipulator with prescribed performance constraints, considering the inertia uncertainties, internal disturbances and input saturation. An inherently continuous adaptive controller is proposed by incorporating non-singular fixed-time sliding mode control, prescribed performance control (PPC), and auxiliary compensation. First, a modified non-singular fast fixed-time terminal sliding surface is constructed, which has a shorten convergence time than the conventional fixed-time sliding surface. Unlike the existing complicated PPCs, a simple structure controller is developed to satisfy prescribed performance constraints through a unique tangent-type PPC technique. The input saturation is then compensated adaptively by an auxiliary mechanism. The Lyapunov theory thoroughly validates the stability and fixed-time convergence of the closed-loop tracking system. With the suggested control scheme, the system states can converge quickly to a small neighbourhood around the origin within a preassigned time, while the position tracking error can be confined within a prescribed performance bounds even in the presence of input saturation. Compared to the existing tracking methods, the suggested control approach has the advantages of faster transient convergence, higher steady-state tracking precision, and stronger robustness. Simulation comparisons demonstrate the effectiveness and superiority of the proposed controller.

Global air traffic demand has shown rapid growth for the last three decades. This growth led to more delays and congestion within terminal manoeuvring areas (TMAs) around major airports. The efficient use of airport capacities through the careful planning of air traffic flows is imperative to overcome these problems. In this study, a mixed-integer nonlinear programming (MINLP) model with a multi-objective approach was developed to solve the aircraft sequencing and scheduling problem for mixed runway operations within the TMAs. The model contains fuel cost functions based on airspeed, altitude, bank angle, and the aerodynamic characteristics of the aircraft. The optimisation problem was solved by using the $\varepsilon$-constraint method where total delay and total fuel functions were simultaneously optimised. We tested the model with different scenarios generated based on the real traffic data of Istanbul Sabiha Gökçen Airport. The results revealed that the average total delay and average total fuel were reduced by 26.4% and 6.7%, respectively.

Airborne particles, such as dust and volcanic ash, pose a serious hazard to aircraft in flight due to their potential to cause erosion damage to engine components. It is crucial to anticipate and address the impact of erosion wear on engine performance and safety. This study aims to enhance our understanding of how volcanic ash particles behave when ingested through a high bypass turbofan engine (HBTFE) and assess the development of erosion wear in the front components. The effects of four different ash samples are assessed in various scenarios of encountering volcanic ash during cruise flight conditions. First, the flow solution is obtained for all front components, including the Pitot intake, spinner, fan, inlet guide vanes (IGVs), outlet guide vanes (OGVs), and connecting ducts. Based on the flow data, the particle motion equations are solved step by step using an in-house trajectory and erosion code. This latter adopts the Lagrangian approach, which incorporates a particle-eddy interaction model and includes probabilistic descriptions for the release positions of particles, sizes, and restitution factors. The finite element method (FEM) is used to track particles through the computational cells and determine impact positions and conditions. As a result, the Pitot intake design seems to prevent many ash particles from reaching the fan blade beyond 80% of the span. The fan blade leading edge (LE) exhibits extreme erosion on both sides. The blade’s pressure side (PS) displays erosion spreading practically on the entirety of the surface, especially near the trailing edge (TE). In contrast, the suction side (SS) has scattered erosion at lower rates. Furthermore, the rotor’s hub presents almost uniform erosion patterns, whereas the shroud depicts scattered erosion. This large fan appears to function as a separator, expelling a significant amount of ash particles through the secondary duct, thereby reducing the engine core’s susceptibility to erosion. Out of the four volcanic ash samples, those from the Kelud and Etna volcanoes appear to cause the highest hourly eroded mass, about twice as much as the samples from the Chaiten and Eyjafjallajokull volcanoes.

In this study, fretting wear in riveted lap joints of aluminium alloy plates was investigated. For the fretting test, 6061-T6 aluminium alloy plates, which are widely used in aircraft construction, and blind rivets were used. Experiments were carried out using a computer controlled Instron testing machine. Fretting surface roughness, microhardness was investigated by metallographic techniques and scanning electron microscopy.

Tensile load cycles in the riveted lap joint were found to cause damage to all surfaces. Two contact surfaces where friction occurs were investigated. The contact surface of the lower plate with the upper plate, and the contact surface of the lower plate with the rivet head. As load and cycles increased, fretting scars and surface roughness increased.

Consequently, it has been determined that fretting damages occur between the contact surface of the plates and between the plate and the rivet contact surface.

The variability in ground manoeuvre occurrences for aircraft landing gear is intrinsically linked to the airport geometries served by aircraft in-service and consequently, the cyclic loads that landing gear carry are driven by the route network and characteristics of aircraft operators. Currently, assumptions must be made when deriving fatigue load spectra for aircraft landing gear, which may fail to capture the operator characteristics, potentially leading to design conservatism. This paper presents the enhanced characterisation of ground turning manoeuvres within the Automatic Dependent Surveillance-Broadcast (ADS-B) trajectories for six narrow-body aircraft across a full-service carrier (FSC) and a low-cost carrier (LCC) fleet. The methodology presented within this paper employs ADS-B latitude and longitude information to overcome limitations of previous approaches, increasing the rate of correct manoeuvre identification within ADS-B trajectories to 77% of flights from the 50% rate achieved previously. When characterising the ground manoeuvres across 3,000 flights, significant differences in manoeuvre occurrences were observed between individual aircraft within the LCC fleet and between the FSC and LCC fleets. The occurrence of tight and pivot turns were shown to vary across the six aircraft with six and eight fatigue-critical turns being performed by the FSC and LCC fleet for every 10 flights performed. In addition, it was observed that the direction of fatigue critical turns is biased in specific directions, suggesting that individual main landing gear assemblies will accumulate fatigue damage at an increased rate, leading to greater justification for operator-specific spectra and structural health monitoring of aircraft landing gear.

Airspace control plays an important role in the safety and fluidity of air traffic. A fundamental service for this purpose is audio communication through frequencies in the VHF bands. This paper describes the evaluation of the audio degradation of voice transmissions from control centre to the aircraft. The effects of more than one station broadcasting on the same frequency with carrier offset (climax mode) are analysed using perceptual evaluation of speech quality (PESQ) perceptual model. Comparative studies are performed to verify the degree of degradation of different audio transport systems and climax situation.

The traditional ground collision avoidance system (GCAS) makes avoidance decisions based on predicted collision time, without considering the impact of terrain environment and dynamic changes in load factor on avoidance decisions. This increases the risk of ground collisions for the aircraft. To solve the problem, a GCAS with multi-trajectory risk assessment and decision function is proposed. Firstly, a variety of predicted flight avoidance trajectories are established within the final manoeuvering capability of the aircraft. Secondly, for each predicted trajectory, the uncertain length between adjacent prediction points is used to construct a rectangular distance bin, and the terrain data below the avoided trajectory is extracted. Finally, the regret theory is used to establish a multi-attribute avoidance decision model to evaluate and prioritise the risk of collision avoidance trajectories, to provide effective collision avoidance decision for pilots. The algorithm is tested and verified with real digital elevation model and simulated flight data, and compared with traditional GCAS. Simulation results show that the proposed algorithm can comprehensively consider manoeuvering performance and threatening terrain, and provide safe and effective avoidance decisions for pilots.

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.

Reynolds Averaged Navier-Stokes (RANS) simulations are performed to investigate the aerodynamic characteristics of the NASA Common Research Model (CRM) in its high-lift (HL) configuration in close proximity to the ground. The RANS simulations are conducted at a moderate Reynolds number of $Re = 5.49 \times {10^6}$ and $M = 0.2$ with the use of the Spalart-Allmaras (SA) turbulence model. out of ground effect (OGE) simulation results are validated against available wind tunnel data before proceeding to in ground effect (IGE) simulations. The obtained computational results in the immediate vicinity of the ground with asymmetric aircraft attitudes demonstrate significant changes in the longitudinal and lateral-directional aerodynamic characteristics, which should be taken into account in flight dynamics analysis of aircraft during take-off and landing in crosswind conditions.

The control of a wing-in-ground craft (WIG) usually allows for many needs, like cruising, speed, survival and stealth. Various degrees of emphasis on these requirements result in different trajectories, but there has not been a way of integrating and quantifying them yet. Moreover, most previous studies on other vehicles’ multi-objective trajectory is planned globally, lacking for local planning. For the multi-objective trajectory planning of WIGs, this paper proposes a multi-objective function in a polynomial form, in which each item represents an independent requirement and is adjusted by a linear or exponential weight. It uses the magnitude of weights to demonstrate how much attention is paid relatively to the corresponding demand. Trajectories of a virtual WIG model above the wave trough terrain are planned using reward shaping based on the introduced multi-objective function and deep reinforcement learning (DRL). Two conditions are considered globally and locally: a single scheme of weights is assigned to the whole environment, and two different schemes of weights are assigned to the two parts of the environment. Effectiveness of the multi-object reward function is analysed from the local and global perspectives. The reward function provides WIGs with a universal framework for adjusting the magnitude of weights, to meet different degrees of requirements on cruising, speed, stealth and survival, and helps WIGs guide an expected trajectory in engineering.

A high-load counter-rotating compressor is optimised based on the method of coupling aerodynamic optimisation technology and computational fluid dynamics, and the flow structures in the passage are analysed and evaluated by vorticity dynamics diagnosis. The results show that the aerodynamic performance of optimised compressor are obviously improved at both design point and off-design point. By comparing the distribution characteristics of vorticity dynamics parameters on the blade surface before and after the optimisation, it is found that BVF (boundary vorticity flux) and circumferential vorticity can effectively capture high flow loss regions such as shock waves and secondary flow in the passage. In addition, the BEF (Boundary enstrophy flux) diagnosis method based on the theory of boundary enstrophy flux is developed, which expands the application scenario of the boundary vorticity dynamics diagnosis method. The change of vorticity dynamics parameters shows blade geometric parameters’ influence on the passage’s viscous flow field, which provides a theoretical basis for the aerodynamic optimisation design.

Runway overruns (ROs) are the result of an aircraft rolling beyond the end of a runway, which is one of the accident’s types that most frequently occurs on aviation. The risk of an RO arises from the synergistic effect among its precursors, such as unstable approaches, long touchdowns and inadequate use of deceleration devices. To analyse this complex socio-technical system, the current work proposes a customised functional resonance analysis method, called FRAM-FDM, as traditional techniques of risk and safety assessment do not identify the interactions and couplings between the various functional aspects of the system itself, especially regarding human and organisational components. Basically, FRAM-FDM is the coupling of a traditional FRAM with flight data monitoring (FDM) techniques, used here to quantify the variabilities of the flight crew performance while executing the required activity (i.e. the landing). In this proposal, these variabilities (i.e. the FRAM functions aspects) are aggregated by the addend of a logistic regression, resulting in a model to evaluate the flare operations and the brake application profile effect on the remaining distance to the end of the runway, used as a reference to classify the landing as acceptable or not. The present application of the FRAM-FDM assesses the operational risk of a sample fleet in overrunning the runway during landing, highlighting the brake pedal application profile as the most relevant contributor. The model improves the knowledge about the system behaviour, being useful to direct flight crew training.

Accurately evaluating the aerodynamic performance of the missile with damaged structures is very important for the subsequent flight control strategy. At present, few researchers have studied the aerodynamic characteristics of damaged supersonic cruise missiles. Based on CFD (computational fluid dynamics) solutions and the dynamic derivative identification method, the differences in static and dynamic characteristics between the damaged and undamaged models are compared. The results indicate that when the extent of damage increases, the change rate of drag coefficient at larger AoA (angle-of-attack) is greater than that at the smaller AoA. On the contrary, the change rate of lift coefficient at larger AoA is smaller than that at smaller AoA. Meanwhile, the absolute value of the static pitch moment decreases, but the absolute value of the roll moment increases. Damage causes a change in the absolute values of the pitch and roll dynamic derivatives, and the dynamic derivatives do not vary monotonically with the increase of AoA. The turning point occurs at about $\alpha$ = 5°. The areas of the hysteresis loops of the pitch-roll coupling moment increase, which makes the dynamic coupling characteristic between the pitch and roll directions increase. Finally, the maximum allowable damage extent of the missile wing that can achieve static trim is obtained and validated by controlling the deflection of the four rudders.

Due to the safety threats caused by icing, the de-icing system is essential in the aviation industry. As an effective method, the electromechanical de-icing system (EDS) is a new ice-protection system based on mechanical vibration principles. For the majority of the current research on system de-icing capability estimation, the effect of impedance-matching is not considered. Impedance matching plays a very important role in improving the performance of the electromechanical system, so we must also consider the impact of impedance matching when designing the EDS. In the present study, a de-icing capability prediction method considering the impact of an impedance-matching device is established based on experimental and numerical methods. The results indicate that the impedance-matching effect has no impact on the mechanical vibration of the structure for the same load power. Meanwhile, impedance-matching devices can significantly improve the power factor and increase the interface shear stress/strain for de-icing. Eight different vibrational modes were tested, and the experimental results showed that the actual interface shear strain after impedance matching is inversely proportional to the de-icing time. The verification experiments were conducted and the accuracy of the proposed prediction method was verified.

The aviation industry’s efforts to reduce carbon emissions have driven the rapid development and scale-up of sustainable aviation fuels (SAFs). SAFs have the potential to significantly reduce CO2 lifecycle emissions by up to 80% in comparison to Jet A and other conventional fossil-derived jet fuels. For multiple logistical and practical reasons, it is preferable to ensure that SAFs are ‘essentially identical’ (also referred to as ‘drop-in SAF’) to conventional jet fuel in terms of their performance, durability and compatibility with existing hardware systems. Because the majority of SAFs are not identical (non-drop-in) to conventional jet fuel, they have not been approved for use in their neat (100%) form. Instead, these non-identical SAFs are named synthetic blend components (SBC) as they are blended with conventional fuels to different extents per ASTM D7566-23a. It should be noted that there are on-going efforts to develop non-drop in SAF specifications to broaden their proliferation and maximise the aviation industries’ ability to reduce CO2 lifecycle emissions. One very important area of focus is the compatibility of SAFs with engine and fuel system seals, specifically understanding the dynamics of elastomeric seals. To address this, a novel approach has been developed to measure seal dynamics in flowing fuel. This technique has been applied to study the dynamic seal behaviour of four industrially relevant elastomer seals commonly employed in aviation fuel systems. The study involved three test fuels: (i) conventional fossil-derived Jet A, neat hydroprocessed esters and fatty acids (HEFA) SAF, and neat alcohol to jet (ATJ) SAF. Notably, both HEFA and ATJ fuels contain 0% aromatics, in contrast to Jet A, which typically contains around 17% aromatics by volume. The novel fuel-elastomer test rig used in this study was designed to simulate a practical scenario in which fuel flows through the inner surface of a pre-loaded static O-ring. The results of these tests demonstrate that the behaviour of different nitrile elastomers is unique to their formulation, and in all cases, the behaviour in HEFA and ATJ SAF differs significantly from that in Jet A. However, new fuel approval tests may only list one type of elastomer for evaluation, for example the ‘Fit-for-Purpose’ test in ASTM D4054-22 Tier 2 lists one specific nitrile. The findings of this study highlight the complexities of fuel-elastomer interactions within nominally identical chemical families and emphasise the potential risks of assessing compatibility based on tests conducted with a single member of a chemical family.

In the present work, a folding wing system (FWS) was developed for guided ammunitions, so that the swept-back angle could be adjusted during both gliding and diving phases. Unlike previous designs, the FWS does not have any fixing mechanisms or brake elements, and it provides folding functionality to reduce the drag force during the terminal phase. We conducted mechanism design, manufactured the FWS, performed system identification and designed various controllers including linear quadratic regulator (LQR), linear quadratic integrator (LQI), sliding mode control (SMC) and second-order sliding mode control (SOSMC) to adjust and hold the desired swept-back angles. Then, the performance of the FWS was tested experimentally under two different flight scenarios, with and without aerodynamic loads. While all controllers operated with almost zero steady-state error (SSE) in the absence of aerodynamic loads, the SOSMC was the most effective controller under aerodynamic loads, considering SSE, delay, chattering, and energy consumption.

It is a simple method to identify the hub dynamic loads of rotor by measuring the vibration responses on helicopter fuselage. However, the identification accuracy of the hub dynamic loads is related to the layout or placement of measuring points on the fuselage. The identification will be inaccurate if the layout of measuring points on the fuselage is unreasonable to result in the “ill-conditioned” frequency response function (FRF) matrix measured on the fuselage. In order to avoid the inaccurate identification due to the “ill-conditioned” measured FRF matrix, an accurate method for identifying the hub dynamic loads of rotor by vibration measurement on helicopter fuselage is proposed in this paper. In the proposed method, the reasonable layout of the measuring points on the fuselage for the “well-conditioned” measured FRF matrix can be obtained according to the condition number of the measured FRF matrix on the fuselage, and then the hub dynamic loads of rotor can be accurately identified. The simulation and experiment of the identification of the hub dynamic loads on a dynamically similar frame structure of a helicopter cockpit floor have verified the effectiveness and accuracy of the proposed method.

Sandwich composites are widely used in aerospace materials thanks to their low weight and high strength properties. The purpose of this study is to observe the effects of polyurethane foam filling on honeycomb core structures produced by additive manufacturing in terms of mechanical strength and moisture absorption properties. Within the scope of the study, honeycomb structures were produced by a 3D printer using polylactic acid (PLA) filament. Then, the honeycomb core was filled with polyurethane foam, which is supplied in liquid form. After the core material was given its final form, it was combined with an epoxy and carbon fibre facesheet material using the vacuum infusion technique. After the sandwich composite production was completed, in-plane compression, three-point bending, shear, and moisture absorption tests were applied. The polyurethane foam filling greatly increased the mechanical strength, but slightly more moisture absorption occurred in this structure compared to a hollow honeycomb structure.

There are many deviation sources in the assembly process of aircraft complex thin-walled structures. To get important factors that affect quality, it is crucial to diagnose the key deviation resources. The deviation transfer between deviation sources and assembly parts has the characteristics of small sample size, nonlinearity, and strong coupling, so it is difficult to diagnose the key deviation sources by constructing assembly dimension chains. Therefore, based on the deviation detection data, transfer entropy and complex network theory are introduced. Integrating the depth-first traversal algorithm with degree centrality theory, a key deviation diagnosis method for complex thin-walled structures is proposed based on weighted transfer entropy and complex networks. The application shows that key deviation sources that affect assembly quality can be accurately identified by the key deviation source diagnosis method based on complex networks and weighted transfer entropy.