In tandem with the fast-growing demand for Unmanned Aerial Vehicles (UAVs) for surveillance and reconnaissance, advanced controllers for these critical systems are needed. This paper proposes a flight dynamics controller design that considers various uncertainties for the Hydra Technologies UAS-S4 Ehécatl. In order to be realistic, in addition to flight dynamics nonlinearities, three main sources of uncertainties are considered, as those caused by unknown controller’s parameters, modeling errors, and external disturbances. A Robust adaptive fuzzy logic controller is designed, in charge of nonlinear flight dynamics in presence of a variety of uncertainties. The nonlinear flight dynamics is modeled based on the Takagi-Sugeno method relying on the soft association of local linear models. Since this controller is model-based, an optimal reference model is defined, which is stabilised by the Linear Quadratic Regulator procedure. A fuzzy logic controller is then designed for the nonlinear model. Lastly, with the aim to handle the uncertainties, the gains of the fuzzy controller are reconfigured, and are continuously adjusted by Lyapunov-based robust adaptive laws. The performance of the UAS-S4 Robust adaptive fuzzy logic controller is evaluated in terms of lateral and longitudinal flight dynamics stabilisation, and the reference model state variables tracking under various uncertainties.

The cost of Reynolds-Averaged Navier-Stokes simulations can be restrictive to implement in aeromechanics design and analysis of vertical lift configurations given the cost to resolve the flow on a mesh sufficient to provide accurate aerodynamic and structural loads. Dual-solver hybrid methods have been developed that resolve the configuration and the near field with the Reynolds-Averaged Navier-Stokes solvers, while the wake is resolved with vorticity-preserving methods that are more cost-effective. These dual-solver approaches can be integrated into an organisation’s workflow to bridge the gap between lower-fidelity methods and the expensive Reynolds-Averaged Navier-Stokes when there are complex physics present. This paper provides an overview of different dual-solver hybrid methods, coupling approaches, and future efforts to expand their capabilities in the areas of novel configurations and operations in constrained and turbulent environments.

This paper proposes a novel real-time formation control and obstacle avoidance algorithm for multiple fixed-wing UAVs. A formation control algorithm is designed by a combination of the virtual structure, leader-follower, and artificial potential fields methods and harnessing the advantages of those approaches. The kinematic and dynamic constraints of fixed-wing UAVs are considered in the path planning. The performance of the proposed algorithm is examined through simulation in Matlab software by applying the translational dynamics of fixed-wing UAVs. Simulations of different complex scenarios demonstrate the effectiveness of the presented formation flight algorithm through generating multiple efficient paths, which are fully consistent with the functional constraints of the UAVs in the presence of obstacles.

A spacetime formulation is presented to solve unsteady aerodynamic problems involving large deformation or topological change such as store separation, slat and flap deployment or spoiler deflection. This technique avoids complex CFD meshing methods, such as Chimera, by the use of a finite-volume approach both in space and time, and permits a locally varying real timestep. The use of a central-difference scheme in the time direction can yield non-physical transient solutions as a consequence of information travelling backwards in time. Therefore, an upwind formulation is provided and validated against one-dimensional and two-dimensional test cases. A hybrid formulation (central in space, upwind in time) is also given and unsteady cases are computed for a spoiler and spoiler/flap deployment, with all three formulations compared, demonstrating that the use of an upwind time stencil yields more representative physical solutions and improves the rate of convergence.

Weapon target allocation (WTA) is an effective method to solve the battlefield fire optimisation problem, which plays an important role in intelligent automated decision-making. We researched the multitarget allocation problem to maximise the attack effectiveness when multiple interceptors cooperatively attack multiple ground targets. Firstly, an effective and reasonable fitness function is established, based on the situation between the interceptors and targets, by comprehensively considering the relative range, relative angle, speed, capture probability and radiation source matching performance and thoroughly evaluating them based on the advantage of the attack effectiveness. Secondly, the optimisation performance of the particle swarm optimisation (PSO) algorithm is adaptively improved. We propose an adaptive simulated annealing-particle swarm optimisation (SA-PSO) algorithm by introducing the simulated annealing algorithm into the adaptive PSO algorithm. The proposed algorithm can enhance the convergence speed and overcome the disadvantage of the PSO algorithm easily falling into a local extreme point. Finally, a simulation example is performed in a scenario where ten interceptors cooperate to attack eight ground targets; comparative experiments are conducted between the adaptive SA-PSO algorithm and PSO algorithm. The simulation results indicate that the proposed adaptive SA-PSO algorithm demonstrates great performance in convergence speed and global optimisation capabilities, and a maximised attack effectiveness can be guaranteed.

This paper presents a new approach to force fighting equalisation in a redundant active-active-active rudder actuation system that is used for the primary flight control system of a turboprop regional aircraft. The related coupled problem of force fighting scenario, and the hydraulic architecture of electronic-hydrostatic actuator (EHA) are analysed, the mathematical model of the EHA system is built. The virtual test bench is designed to evaluate the performance of the force fighting equalisation strategy. The proposed methodology is tested on an iron bird test rig. The physical experiment shows that the fighting force is minimised under all flight conditions, meets the low cost requirement and can be a very reliable system. The proposed methodology can be applied to other types of aircraft’ flight actuation systems.

Compression systems of modern, civil aircraft engines consist of three components: Fan, low-pressure compressor (LPC) and high-pressure compressor (HPC). The efficiency of each component has improved over the last decades by means of rising computational power which made high level aerodynamic optimisations possible. Each component has been addressed individually and separated from the effects of upstream and downstream components. But as much time and effort has been spend to improve performance of rotating components, the stationary inter-compressor duct (ICD) has only received minor attention. With the rotating compression components being highly optimised and sophisticated their performance potential is limited. That is why more aggressive, respectively shorter, ICDs get more and more into the focus of research and engine manufacturers. The length reduction offers high weight saving and thus fuel saving potential as a shorter ICD means a reduction in aircraft engine length. This paper aims at evaluating the impact of more aggressive duct geometries on LPC and HPC performance. A multi objective 3D computational fluid dynamics (CFD) aerodynamic optimisation is performed on a preliminary design of a novel two spool compressor rig incorporating four different operating line and two near-stall (NST) conditions which ensure operability throughout the whole compressor operating range. While the ICD is free to change in length, shape and cross-section area, the blades of LPC and HPC are adjusted for changing duct aerodynamics via profile re-staggering to keep number of free parameters low. With this parametrisation length, reductions for the ICD of up to 40% are feasible while keeping the reduction in isentropic efficiency at aerodynamic design point for the compressor below 1%pt. Three geometries of the Pareto front are analysed in detail focusing on ICD secondary flow behaviour and changes of aerodynamics in LPC and HPC. In order to asses changes in stall margin, speedlines for the three geometries are analysed.

For the safety problems caused by the limited landing space of the deck during the arresting process of the carrier-based aircraft, a dynamic model of the carrier-based aircraft’s landing and arresting is built. Based on the batch simulation method, the lateral dynamics safety envelope of the aircraft during the arresting was defined, and the dynamic response of the key points in the envelope during the arresting process was investigated. Subsequently, the influence of engine thrust and aircraft quality on the arresting safety envelope was studied based on reasonable safety evaluation indicators, and the safety status envelope of the deck arresting was given. Then, the particular Hamilton-Jacobi partial differential equation is used to obtain the lateral dynamics safety envelope of the carrier-based aircraft in the process of landing and arresting by backward inversion. Results indicate that engine thrust and landing quality have little effect on the yaw angle in the arresting safety boundary during the arresting. Additionally, with the engine thrust and landing quality increase, the maximum safe off-centre distance gradually decreases, and the safety boundary decreases accordingly. During the phase of landing glide, the engine thrust and quality have little effect on the maximum safe eccentric distance. When the engine thrust is increased by 40%, the maximum safe yaw angle is reduced from 0.3°, and the safety boundary is reduced by 4.2%. When the aircraftquality increases by 40%, the maximum safe yaw angle is reduced by 0.4°, and the safety boundary is reduced by 2.8%. The findings of this paper can provide framework for the research on theaircraft-to-carrier dynamic matching characteristics of the carrier-based system, and is of great significance to the research on improving the safety of the carrier-based aircraft landing arresting.

In recent years, Performance Based Logistics (PBL) has been increasingly used to reduce the product lifecycle cost. As a result, the existing logistics analysis methods need to be reassessed due to the difference in PBL from the classical approach. This study considers the Spare Parts Allocation (SPA) and Level of Repair Analysis (LORA), which are the most commonly used problems within PBL, and are the subject of analysis. A comprehensive, multi-objective simulation-optimisation model is developed for a military aircraft operations case study, with the objectives of minimising total cost and maximising total flight hours. The Design of Experiment (DOE) method is used for determining simulation-optimisation parameters. Technique for Order of Preference by Similarity to Ideal Solution (TOPSIS) method is used for the first time in order to select the best design points. Simulated Annealing (SA) was used for multi-objective optimisation with the SPA model being solved first. Then the LORA problem was added to this model as a decision variable and the effects of the integrated solution (SPA+LORA) were examined. The results show that the integrated solution yields remarkably better results in terms of flight hours and cost when compared to the SPA optimisation approach alone. Moreover, the proposed model provides a profit-centric approach and can be efficiently used as a decision support tool for both customers and suppliers for difficult and complex logistics support activities.

Climate change is increasing global-mean tropospheric temperatures, but the localised trends are uneven, including cooling the lower stratosphere and lifting the tropopause. The wind speeds are also being modified, both at the surface and aloft. A further effect, additional to wind and temperature alone, is of increasing fluctuations and severity of extreme weather. These are impacting air transport, and this will continue. The effects are known to include increased take-off distances where excess runway lengths exist and reduced payloads where they do not, increased en-route flight times, increased frequency and severity of encounters with clear air turbulence in some regions, changed patterns of wildlife — particularly bird — activity in some regions (potentially also for other anthropogenic reasons) are shifting locations of flight safety hazards, and increased burdens upon airport and associated infrastructure. There is increasing understanding and acknowledgment by companies and authorities of these effects and the importance of mitigating them, although this is not universal and there are as yet no universally understood best practices for air transport climate change mitigation.

Variable camber flap technology can adjust the spanwise circulation distribution, thereby reducing the induced drag. Therefore, the concept of variable camber flap is introduced into the design of propeller aircraft wing, and the design for drag reduction of propeller aircraft is carried out. The numerical simulation of the propeller aircraft is carried out by using the actuator disc method with non-uniform distribution of radial and circumferential loads. Through the unsteady simulation of a single propeller, the aerodynamic load on a periodic propeller is extracted as a boundary condition to the steady simulation of the full aircraft. The load extracted by the actuator disc is compared with the unsteady simulation result, which verifies the reliability of the method. The design for drag reduction at cruise and climb design conditions are respectively carried out with the variable camber flap technology. The variable camber cruise configuration is evaluated at both the begin and end cruise conditions. The results show that, after the flaps deflecting at a small angle according to the circulation distribution, the camber distribution of the wing is adjusted to make the circulation distribution closer to the elliptical circulation distribution. At the design cruise condition, the drag coefficient is reduced by 1.4 counts, and the lift-drag ratio increase by 0.1. At both begin and end cruise conditions, the drag coefficient decreases by 1 count, and the lift-drag ratio increases by 0.07. At the design climb condition, the drag coefficient decreases by 1 count, and the lift-to-drag ratio increases by 0.09.

The aim morphing of this study is to detect and reconstruct a fault in angle-of-attack sensor and pitot probes that are components in commercial aircrafts, without false alarm and no need for additional measurements. Real flight data collected from a local airline was used to design the relevant system. Correlation analysis was performed to select the data related to the angle-of-attack and airspeed. Fault detection and reconstruction were carried out by using Adaptive Neural Fuzzy Inference System (ANFIS) and Artificial Neural Networks (ANN), which are machine-learning methods. No false alarm was detected when the fault test following the fault modeling was carried out at 0–1 s range by filtering the residual signal. When the fault was detected, fault reconstruction process was initiated so that system output could be achieved according to estimated sensor data. Instead of using the methods based on hardware redundancy, we designed a new system within the scope of this study.

Fibre-reinforced polymer (FRP) composites generally have a layered architecture and are commonly manufactured with thermosetting resins—making them susceptible to interlaminar fracture (i.e. delamination), which is often a major concern in structurally critical applications. As a result, various approaches have been explored to enhance interlaminar fracture resistance. This review focuses on third-phase toughener inclusions, which offer opportunities to create damage resistant and damage tolerant structures without significantly adding weight or reducing in-plane mechanical properties. These toughener inclusions, typically introduced in the interlaminar regions, are divided into two categories herein: particle fillers and non-woven fibre veils. The advantages and limitations of both types are discussed, and the potential of the two approaches is evaluated using published data, aiming to provide an overview of the current understanding and challenges in designing and manufacturing safe and reliable composite structures.

Prediction of stall before it occurs, or detection of stall is crucial for smooth and lasting operation of fans and compressors. In order to predict the stall, it is necessary to distinguish the operational and stall regions based on certain parameters. Also, it is important to observe the variation of those parameters as the fan transitions towards stall. Experiments were performed on a contra-rotating fan setup under clean inflow conditions, and unsteady pressure data were recorded using seven high-response sensors circumferentially arranged on the casing, near the first rotor leading edge. Windowed Fourier analysis was performed on the pressure data, to identify different regions, as the fan transits from the operational to stall region. Four statistical parameters were identified to characterise the pressure data and reduce the number of data points. K-means clustering was used on these four parameters to algorithmically mark different regions of operation. Results obtained from both the analyses are in agreement with each other, and three distinct regions have been identified. Between the no-activity and stall regions, there is a transition region that spans for a short duration of time characterised by intermittent variation of abstract parameters and excitations of Fourier frequencies. The results were validated with five datasets obtained from similar experiments at different times. All five experiments showed similar trends. Neural Network models were trained on the clustered data to predict the operating region of the machine. These models can be used to develop control systems that can prevent the stalling of the machine.

The tip leakage flow generates a large amount of aerodynamic losses in a zero inlet swirl turbine rotor (ZISTR), which directly uses the axial exit flow downstream of a combustion chamber without any nozzles. To reduce the tip leakage flow loss and improve the efficiency for the ZISTR, a front suction side winglet is employed on the blade tip, and the effect of winglet width is numerically investigated to explore its design space. It is found that, a suction side leading edge horseshoe vortex (SHV) on the blade tip plays a crucial role in mitigating the tip leakage flow loss. This SHV rotates in the reverse direction to the leakage vortex, so it tends to break the formation of the leakage vortex near the front part of suction side. With a larger winglet width, the SHV stays longer time on the blade tip and leaves it at a further downstream location. This increases the time and the contact area of the interaction between the SHV and the leakage vortex, so the leakage vortex is further weakened. Thus, the tip leakage flow loss is reduced, and the efficiency is improved. However, a larger winglet width also increases the heat load of the blade due to a larger blade surface area. The ZISTR designed with the winglet width equal to 2.1% blade pitch achieves a great trade-off between efficiency and heat load that the efficiency is improved by 0.85% at an expense of 1.2% increment of the heat load. Besides, for the blade using this winglet, the mechanical stress due to the centrifugal, aerodynamic and thermal load is acceptable for the engine application. This investigation indicates a great potential in the improvement of efficiency for the ZISTR using a blade tip winglet designed on the front suction side.

A shock-induced separation loss reduction method, using local blade suction surface shape modification (smooth ramp structure) with constant adverse pressure gradient with the consideration of radial equilibrium effect to split a single shock foot into multiple weaker shock wave configuration, is investigated on the NASA Rotor 37 for promoting aerodynamic performance of a transonic compressor rotor. Numerical investigation on baseline blade and improved one with blade modification on suction side has been conducted employing the Reynolds-averaged Navier–Stokes method to reveal flow physics of ramp structure. The results indicate that the passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot generating moderate adverse pressure gradient on ramp profile, which is beneficial for mitigating the shock foot and shrinking flow separation region as well. In addition, the radial secondary flow of low-momentum fluids within boundary layer is decreased significantly in the region of passage shock-wave/boundary-layer interaction on blade suction side, which mitigates the mass flow and mixing intensity of tip leakage flow. With the reduction of flow separation loss induced by passage shock, the adiabatic efficiency and total pressure ratio of improved rotor are superior to baseline model. This study herein implies a potential application of ramp profile in design method of transonic and supersonic compressors.

In real gas turbines, multiple nozzles are used instead of a single-nozzle; therefore, interactions between flames are inevitable. In this study, the effects of flame-flame interaction on the emission characteristics and lean blowout limit were analysed in a CH4-fueled single- and dual-nozzle combustor. OH* chemiluminescence imaging showed that a flame-interacting region, where the two flames from the nozzles were merged, was present in the dual-nozzle combustor, unlike the single-nozzle combustor. Flow-field measurements using particle image velocimetry confirmed that a faster velocity region was formed at the flame merging region, thereby hindering flame stabilisation. In addition, we compared the emission indices of NOx and CO between the two combustors. The emission indices of CO were not significantly different; however, a distinct effect of flame-flame interaction was indicated in NOx. To understand the effect of flame-flame interaction on NOx emissions, we measured temperature distribution using a multi-point thermocouple. Results showed that a wider high-temperature region was formed in the dual-nozzle combustor compared to the single-nozzle combustor; this was attributable to the high OH* chemiluminescence intensity in the flame-interacting region. Furthermore, it was confirmed that the size of this interacting region caused the deformation of the temperature distribution in the combustor, which can induce a difference in the increase ratio of NOx emission between high and low equivalence ratio ranges. In conclusion, we confirmed that flame-flame interaction significantly affected temperature distribution in the downstream of the flame, and the change in temperature distribution contributed primarily to the varying concentration of the emission gas.

A new aerodynamic open-circuit test rig for studying boundary layer ingestion (BLI) propulsion has been developed by National Research Council of Canada. The purpose is to demonstrate the advantages of BLI in reducing the power required for a given thrust and to validate the performance of BLI fan concepts. The rig consists of a boundary layer generator to simulate boundary layer development over an aircraft fuselage. The boundary layer generator can be used to create a natural boundary layer due to skin friction but also comprises an array of perforated plates through which pressurised air can be blown to manipulate the boundary layer thickness. The size of the boundary layer thickness can be controlled upstream of the fan blades. Parametric studies of boundary layer thickness were then feasible. The test calibration was conducted to validate the concept.