Considering the shortcomings of current methods for real-time resolution of two-aircraft flight conflicts, a geometric optimal conflict resolution and recovery method based on the velocity obstacle method for two aircraft and a cooperative conflict resolution method for multiple aircraft are proposed. The conflict type was determined according to the relative position and velocity of the aircraft, and a corresponding conflict mitigation strategy was selected. A resolution manoeuvre and a recovery manoeuvre were performed. On the basis of a two-aircraft conflict resolution model, a multi-aircraft cooperative conflict resolution game was constructed to identify an optimal solution for maximising group welfare. The solution and recovery method is simple and effective, and no new flight conflicts are introduced during track recovery. For multi-aircraft conflict resolution, an equilibrium point that maximises the welfare function of the group was identified, and thus, an optimal strategy for multi-aircraft conflict resolution was obtained.

The simulations and assessment of transient performance of gas turbine engines during the conceptual and preliminary design stage may be conducted ignoring heat soakage and tip clearance variations due to lack of detailed geometrical and structural information. As a result, problems with transient performance stability may not be revealed correctly, and corresponding design iterations would be necessary and costly when those problems are revealed at a detailed design stage. To make an engine design more cost and time effective, it has become important to require better transient performance simulations during the conceptual and preliminary design stage considering all key impact factors such as fuel control schedule, rotor dynamics, inter-component volume effect as well as heat soakage and tip clearance variation effects. In this research, a novel transient performance simulation approach with generically simplified heat soakage and tip clearance models for major gas path components of gas turbine engines including compressors, turbines and combustors has been developed to support more realistic transient performance simulations of gas turbine engines at conceptual and preliminary design stages. Such heat soakage and tip clearance models only require thermodynamic design parameters as input, which is normally available during such design stages. The models have been implemented into in-house transient performance simulation software and applied to a model twin-spool turbojet engine to test their effectiveness. Comparisons between transient performance simulated with and without the heat soakage and tip clearance effects demonstrate that the results are promising. Although the introduced heat soakage and tip clearance models may not be as accurate as that using detailed component geometrical information, it is able to include the major heat soakage and tip clearance effects and make the transient performance simulations and analysis more realistic during conceptual and preliminary engine design stage.

In order to intercept a highly manoeuvering target with an ideal impact angle in the three-dimensional space, this paper promises to probe into the problem of three-dimensional terminal guidance. With the goal of the highly target acceleration and short terminal guidance time, a guidance law, based on the advanced fast non-singular terminal sliding mode theory, is designed to quickly converge the line-of-sight (LOS) angle and the LOS angular rate within a finite time. In the design process, the target acceleration is regarded as an unknown boundary external disturbance of the guidance system, and the RBF neural network is used to estimate it. In order to improve the estimation accuracy of RBF neural network and accelerate its convergence, the parameters of RBF neural network are adjusted online in real time. At the same time, an adaptive law is designed to compensate the estimation error of the RBF neural network, which improves the convergence speed of the guidance system. Theoretical analysis demonstrates that the state and the sliding manifold of the guidance system converge in finite time. According to Lyapunov theory, the stability of the system can be guaranteed by online adjusting the parameters of RBF neural network and adaptive parameters. The numerical simulation results verify the effectiveness and superiority of the proposed guidance law.

This paper investigates the feasibility of improving the aircraft landing performance by design the damping orifice parameters of the landing gear using lattice Boltzmann method coupled with the response surface method. The LBM is utilised to simulate characteristics of the damping orifice after model validation. The numerical model of the landing gear using simulated damping force is validated by single landing gear drop test. Based on the numerical model and the response surface functions, the sensitivity analysis and the optimisation design are performed. The maximum error of mean velocity simulated using LBM with experimental data is 7.07% for sharp-edged orifices. Moreover, the numerical model predicts the landing responses adequately, the maximum error with drop test data is 2.51%. The max overloading of the aircraft decreases by 5.44% after optimisation, which proves that this method is feasible to design the damping orifice for good landing performance.

With an annual growth in travel demand of about 5% globally, managing the environmental impact is a challenge. In 2019, the International Civil Aviation Organisation (ICAO) issued emission reduction targets, including well-to-wake greenhouse gas (GHG) emissions reduced at least 50% from 2005 levels by 2050. This discusses several technologies from an aircraft design perspective that can contribute to achieving these targets. One thing is certain: aircraft will look different in the future. The Transonic Truss-Braced Wing and Flying V configurations are promising significant efficiency improvements over conventional configurations. Electric propulsion, in various architectures, is becoming a feasible option for general aviation and commuter aircraft. It will be a growing field of aviation with zero-emissions flight and opportunities for special missions. Lastly, this paper discusses methods and design processes that include all relevant disciplines to ensure that the aircraft is optimised as a complete system. While empirical methods are essential for initial design, Multidisciplinary Design Optimisation (MDO) incorporates models and simulations integrated in an optimisation environment to capture critical trade-offs. Concurrent design places domain experts in one site to facilitate collaboration, interaction, and joint decision-making, and to ensure all disciplines are equally considered. It is supported by a Collaborative Design Facility (CDF), an information technology facility with connected hardware and software tools for design analysis.

Estimation of the structural weight of an Unmanned Combat Aerial Vehicle (UCAV) during conceptual design has proven to be a significant challenge mainly due to its unconventional configuration. We investigate development of a customised approach for structural weight estimation of UCAV based on statistical weight of the manned fighter’s components by applying minor modifications on weight formulations of fuselage, wing, empennage, power plant and landing gear. The modifications are applied by considering the corresponding differences between manned fighters and UCAVs such as manned requirements and mission variances. Some new empirical formulas for estimating the weight of UCAV’s components are proposed. Results for the empty weight estimation are validated against actual values of some well-known UCAVs. Moreover, the structural weight is validated against the benchmark UCAV case studies. The results show that the ratio of structural to takeoff weight for UCAVs is approximately between 20% to 10%. Finally, a generalised equation is developed for estimating the structural weight of UCAVs in conceptual design phase.

This paper proposes an online entry trajectory planning based on predictor corrector algorithm that satisfies terminal constraints, path constraints and geographic constraints for lifting-body entry vehicles. The vehicle is considered as a 3DOF point mass. A piecewise linear bank profile in the altitude-versus-velocity (H-V) plane is used to project the longitudinal trajectory and a heading angle corridor-based bank angle reversal logic is designed to satisfy the geographic constraints simultaneously. The algorithm in longitudinal plane solves the problem of the mismatch between terminal range and height and the algorithm in lateral plane satisfies the no-fly zone constraint. Simulation results with the CAV-H model show that the proposed algorithm can generate entry trajectories satisfying multiple constraints and has certain robustness.

To investigate the downstream rim seal gas ingestion characteristics of a 1.5-stage turbine, the URANS equations were solved numerically using the SST turbulence model. The effects of different purge flow rates and the second vane on the ingestion characteristics of the aft cavity and the nonuniform fluctuations of the main gas path pressure are analysed. The results showed that the aft cavity is affected by the combined effects of the blade and the second vane, and the potential field at the leading edge of the second vane greatly influence the airflow variation in the aft cavity, which enhances the ingress of the mainstream into the wheel-space. The front purge flow weakens the egress between the suction side of the blade and the suction side of the second vane. The potential field at the leading edge of the second vane suppresses the nonuniform distribution of airflow in the aft cavity caused by the rotational effect of the blade.

This study presents a stochastic mixed-integer linear programming model for the aircraft sequencing and scheduling problem. The proposed model aims to minimise the average fuel consumption per aircraft in the Terminal Manoeuvring Area while considering uncertain flight durations for each flight. The tabu search algorithm was selected to solve the problem. The stochastic solution and deterministic solution results were compared to show the benefits of the stochastic solution. The average sample approximation technique was applied to this problem, and enhancement rates of the average fuel consumption per aircraft were 8.78% and 9.11% comparing the deterministic approach

A design of a sub-scale Boundary Layer Ingestion (BLI) fan for a transonic test rig is presented. The fan is intended to be used in flow conditions with varying distortion patterns representative of a BLI application on an aircraft. The sub-scale fan design is based on a design study of a full-scale fan for a BLI demonstration project for a Fokker 100 aircraft. CFD results from the full-scale fan design and the ingested distortion pattern from CFD analyses of the whole aircraft are used as inputs for this study. The sub-scale fan is designed to have similar performance characteristics to the full-scale fan within the capabilities of the test facility. The available geometric rig envelope in the test facility necessitates a reduction in geometric scale and consideration of the operating conditions. Fan blades and vanes are re-designed for these conditions in order to mitigate the effects of the scaling. The effects of reduced size, increased relative tip clearance and thicknesses of the blades and vanes are evaluated as part of the step-by-step adaption of the design to the sub-scale conditions. Finally, the installation effects in the rig are simulated including important effects of the by-pass flow on the running characteristics and the need to control the effective fan nozzle area in order to cover the available fan operating range. The predicted operating behaviour of the fan as installed in the coming transonic test rig gives strong indication that the sub-scale fan tests will be successful.

The aim of this paper is to provide preliminary results on a traffic coordination framework based on stochastic task allocation. General trends and the predicted advent of personal aerial vehicles increase traffic rapidly, but current air traffic management methods admittedly cannot scale appropriately. A hierarchical system is proposed to overcome the problem, the middle layer of which is elaborated in this paper. This layer aims to enable stochastic control of traffic behaviour using a single parameter, which is achieved by applying distributed stochastic task allocation. The task allocation algorithm is used to allocate speeds to vehicles in a scalable way. By regulating the speed distribution of vehicles the conflict rates remain manageable. Multi-agent simulation results show that it is possible to control ensemble dynamics and together with that traffic safety and throughput via a single parameter. Using transient simulations the dynamic performance of the system is analysed. It is shown that the traffic conflict reduction problem can be transformed into a control design problem. The performance of a simple controller is also evaluated. It was shown that by applying the controller, quicker transients can be achieved for the mean speed of the system.

This paper investigates the hydrodynamic near-field of a NACA 16-616 aerofoil over a range of angles-of-attack, encompassing the pre-stall, stall and post-stall flow regimes. In both the static pressure and the pressure fluctuation results, it is shown that each flow regime is easily distinguished, and it is further shown that each regime has different spectral behaviour and boundary layer characteristics. It is found that the NACA 16-616 aerofoil stalls by an abrupt leading-edge mechanism, characterised by a sudden change in the static pressure and unsteady surface pressure spectra between $16^\circ $ and $17^\circ $ angles-of-attack, but of more interest is that there is a secondary yet significant trailing-edge flow separation mechanism occurring upstream of the trailing-edge and moving further upstream as the angle-of-attack increases in the pre-stall regime. A comparison is made between the spectra and coherence of the unsteady surface pressure of the NACA 16-616 aerofoil and those of the classic NACA 0012 aerofoil and shows that such a secondary mechanism has a significant impact for large pre-stall angles-of-attack on the unsteady surface pressure. This will have a significant impact on the radiated far-field sound, distinguishing the NACA 16-616 aerofoil from aerofoils that do not have this secondary mechanism. The existence and extent of this secondary trailing-edge separation mechanism is further shown by the hot-wire anemometry boundary layer velocity results that indicate separation within the pre-stall regime.

Strategic airlift is a crucial capability for any country that wants to protect its global interests around the world. Air refueling may offer more agile and efficient airlift capabilities by increasing cargo aircraft payload and shortening airlift time. We investigated whether air refueling can shorten the total time of an airlift operation and decrease the number of cargo aircraft sorties required in a deployment scenario, especially where the distance between origin and destination is within the range of cargo aircraft. We introduced two mathematical models to compare the total airlift time and number of cargo aircraft required for given origin-destination and tanker base locations and total freight to be moved. We optimised initial cargo and fuel amount for cargo aircraft along with rendezvous point coordinates to minimise total airlift time. We used a numerical example to show that substantial airlift time and cargo aircraft sortie savings are possible through air refueling.

This article focuses on the aerodynamic design of a morphing aerofoil at cruise conditions using computational fluid dynamics (CFD). The morphing aerofoil has been analysed at a Mach number of 0.8 and Reynolds number of $3 \times 10^{6}$ , which represents the transonic cruise speed of a commercial aircraft. In this research, the NACA0012 aerofoil has been identified as the baseline aerofoil where the analysis has been performed under steady conditions at a range of angles of attack between $0^{^{\kern1pt\circ}}$ and $3.86^{^{\kern1pt\circ}}$ . The performance of the baseline case has been compared to the morphing aerofoil for different morphing deflections ( $w_{te}/c = [0.005 - 0.1]$ ) and start of the morphing locations ( $x_{s}/c = [0.65 - 0.80]$ ). Further, the location of the shock wave on the upper surface has also been investigated due to concerns about the structural integrity of the morphing part of the aerofoil. Based upon this investigation, a most favourable morphed geometry has been presented that offers both, a significant increase in the lift-to-drag ratio against its un-morphed counterpart and has a shock location upstream of the start of the morphing part.

A smart morphing winglet driven by piezoelectric Macro Fiber Composite (MFC) is designed to adjust cant angle autonomously for various flight conditions. The smart morphing winglet is composed of the MFC actuator, DC-DC converter, power supply, winglet part and wing part. A hinge is designed to transfer the bending deformation of intelligent MFC bending actuator to rotation of the winglet structure so as to achieve the adaptive cant angle. Experimental and numerical work are conducted to evaluate the performance of smart morphing winglet. It is demonstrated that the proposed intelligent MFC bending actuator has an excellent bending performance and load resistance. This smart morphing winglet exhibits the excellent characteristic of flexibility on large deformation and lightweight. Moreover, a series of wind tunnel tests are performed, which demonstrate that the winglet driven by intelligent MFC bending actuator produces sufficient deformation in various wind speed. At high wind speed, the cant angle of the winglet can reach 16 degrees, which is still considered to be very useful for improving the aerodynamic performance of the aircraft. The aerodynamic characteristics are investigated by wind tunnel tests with various attack angles. As a result, when the morphing winglet is actuated, the lift-to-drag ratio could vary up to 11.9% and 6.4%, respectively, under wind speeds of 5.4 and 8.5m/s. Meanwhile, different flight phases such as take-off, cruise and landing are considered to improve aerodynamic performance by adjusting the cant angle of winglet. The smart morphing winglet varies the aerofoil autonomously by controlling the low winglet device input voltage to remain optimal aerodynamic performance during the flight process. It demonstrates the feasibility of piezoelectric composites driving intelligent aircraft.

This study presents a methodology to estimate the battery consumption of an electric powerplant, based on brushless motors, typically used in light unmanned aerial systems. The methodology models brushless motors through an equivalent circuit obtained from their dynamic behaviour. Propellers’ data are taken from an experimental database. Furthermore, a variable speed controller efficiency is considered in the methodology. All the parameters involved in the model are adjusted by minimising the mean quadratic error of measurements taken in both direct and alternating currents. This model allows designers to predict energy consumption, also if any element of the powerplant changes, such as battery or propeller. Thus, it is useful for selecting the best powerplant for an actual RPAS operation. The results obtained to predict the current consumption of several electric powerplants show a coefficient of determination higher than 0.96. Finally, the methodology is validated by means of a case study of an actual RPAS, where the best powerplant is selected in terms of endurance.

In centrifugal compressors, the identification of flow instability signals from experiments is a difficult problem owing to the nonlinear and non-stationary characteristics. Otherwise, the complicated asymmetric structure of the volute brings a huge challenge to the evolution and circumferential nonuniformity characteristics of the flow instabilities. This paper presents experimental and numerical investigations on internal flow field to understand the flow instability characteristics in a centrifugal compressor. Considering nonlinear and non-stationary signals, a method based on Fourier-transform and variational mode decomposition was introduced to analyse the flow instability characteristics. The Fourier spectrum results show that at 0.21kg/s of 80krpm, the pressure signal has a noticeable high-frequency fluctuation, which indicates that the compressor enters the flow instability state. The variational mode decomposition results show that before a surge, the compressor experiences different flow instability stages: the RI stage, the coexistence stage of RI and stall, and the stall stage. Moreover, obvious circumferential nonuniformity characteristics of flow instabilities were observed during the throttling process. RI first occurred at the 180° circumferential position and then the stall first appeared in the circumferential range of 60° to 240°. The simulation results that it is because that the asymmetric volute causes the adverse pressure gradient inside the impeller passage and a high-pressure region (120°–240°) at the upstream of the impeller inlet. Under this combined action of the two, the effect region of tip leakage vortex expands the upstream of the impeller inlet. Meanwhile, the tip leakage vortex core migrates to a lower span of blades. This study demonstrates the ability to analyse nonlinear and non-stationary signals from a centrifugal compressor via variational mode decomposition, and provides a useful guidance for the identification of flow instability signals.

Wind tunnels remain an essential element in the design and development of flight vehicles. However, graduates in aerospace engineering tend to have had little exposure to the demands of industrial experimental work, particularly at high speed, a situation exacerbated by a lack of up-to-date reference material. In an attempt to fill this gap, this paper presents an overview of the current and near-term status and usage of transonic industrial wind tunnels. The review is aimed at recent entrants to the field, with the aim of helping them make the step from research projects in small university facilities to commercial projects in large industrial facilities. In addition, a picture has emerged from the review that contradicts received wisdom that the wind tunnel is in decline. Globally, the industrial transonic wind tunnel is undergoing somewhat of a renaissance. Numbers are increasing, investment levels are rising, capabilities are being enhanced, and facilities are busy.

A new method for predicting manoeuvre loads on a large transport aircraft with a swept-back wing and a load alleviation system based on control surface deflections is developed. For this purpose, three-dimensional Reynolds-averaged Navier–Stokes (RANS) simulations of the rigid wing–fuselage configuration are performed while the aerodynamics of the tailplane are estimated by means of handbook methods. For a closer analysis, different quasi-steady pitching manoeuvres are chosen based on the CS-25 regulations. One of these manoeuvres is also simulated with active load alleviation, leading to a reduction in the wing-root bending moment by more than 40%. Besides demonstrating the potential of the considered load alleviation system, it is shown which manoeuvres are especially critical in this context and which secondary effects come along with load alleviation.

Pilot-Induced Oscillation (PIO), although an old issue, still poses a significant threat to aviation safety. The introduction of new systems in modern aircraft modifies the human–machine interaction and makes it necessary for research to revisit the subject from time to time. Given the need of aircraft manufacturers to constantly perform PIO tests, this study analysed the feasibility of using three different computational pilot models (Tustin, Crossover and Precision) to simulate PIO conditions. Three aircraft models with different levels of propensity to PIO (original, low propensity and high propensity) were tested, as well as two pilot gain conditions (normal and high). Data were collected for a purely longitudinal synthetic task through simulations conducted in MATLAB®. PIO conditions were detect using a tuned PIO detection algorithm (ROVER). Data were analysed in terms of both whether the pilot models triggered a PIO condition and for how long the condition was sustained. The results indicated that the three pilot models only provoked PIO conditions when high gain inputs were applied. Additionally, Crossover was the only pilot model to trigger a PIO for the three aircraft models. There were also significant differences between the pilot models in the total PIO time, as the Tustin model typically sustained the oscillatory condition for longer.