With the increase of air transportation, some crossing waypoints (CWPs) are becoming bottlenecks in the operation of air traffic networks. This paper presents a CWP operation optimisation framework based on a two-stage optimisation method. First, we considered the interests of airlines and air traffic controllers and established a flight-level dynamic allocation model for the CWP to minimise the flight-level deviation and the number of flight conflicts. A multi-objective, self-adaptive differential evolution-local search hybrid algorithm was used to solve the model in a parallel computing manner. Subsequently, a flight conflict resolution algorithm based on the Monte-Carlo tree search was designed for flight conflicts that existed after the optimisation. Finally, based on real operation data, four experimental scenarios were constructed, and the air traffic operation simulation system was used for experimental validation. For daily traffic and 1.2 times peak traffic scenarios, the average flight-level deviation reduction rates after optimisation were 53% and 39%, and the successful flight conflict resolution rates reached 89% and 75%, respectively. The experimental results showed that this optimisation framework can effectively balance the number of flight conflicts with the efficiency of flight-level usage and directly improve the capacity of the CWP, which can be used as a reference for air traffic control auxiliary decision support systems.

Hypersonic air-breathing propulsion can improve cost and flexibility of Low Earth Orbit (LEO) satellite launch missions. However, at the high flight Mach numbers required for access-to-space, performance margins are extremely tight. Techniques to improve mixing efficiency can push this technology forward. However, these are required to produce a minimal increase in losses and heat loads to be viable. The use of inlet-generated vortices in scramjets for mixing enhancement was previously studied. These vortices interact with the injected fuel plume, stretching it and increasing its effective surface for mixing. Moreover, these vortices are intrinsic to the flowfield. Therefore, contrary to other methods, when using inlet vortices mixing is enhanced without producing additional heat loads or losses. This work studies the vortex-injection interaction through numerical RANS simulations. A non-dimensional variable defining the quality of the plume shape for mixing purposes is proposed. This parameter is used to assess the effect of vortex intensity and injector location on fuel plume shape. The results show the ability of inlet vortices to modify fuel plume shape significantly increasing fuel mixing rate with minimal impact on losses.

The rotating instability in a contra-rotating axial flow compressor is investigated by experiments. Twenty-four pressure sensors were installed on the casing to capture the unsteady flow in the rotor tip region simultaneously. A double-phase-locking technique suitable for the contra-rotating compressor was proposed to characterise the static pressure contours of the rotor tip. The mean and root-mean-square pressure contours indicate that rotating instability occurs before the rotating stall happened, and the rotor tip clearance vortex is located upstream of the rear rotor leading edge plane before stall. Fourier spectrum shows that rotating instability and rotating stall both happened under the stall condition, and the frequency band of rotating instability does not change with the flow rate. In the front rotor, the frequency of rotating instability is half of the blade passing frequency. It is verified that the modal estimation method can be implemented by using the average azimuthal phase velocity, which significantly reduced the number of pressure sensors required. Modal estimation results show that each peak of the rotating instability frequency band corresponds to a unique dominant circumferential mode. By optimising average azimuthal phase velocity, an improved modal estimation method is obtained, which can further improve the reliability of the modal estimation results.

The present study employs the direct simulation Monte Carlo (DSMC) technique to analyse the flow over a cavity, a commonly observed anomaly on a re-entry vehicle’s surface. The flow characteristics are examined for different Mach numbers ( ${\rm{Ma}}$ ) and Knudsen number ( ${\rm{Kn}}$ ). The Mach numbers varied from 5 to 25, while the Knudsen numbers varied from 0.05 to 21.10. The influence of the ${\rm{Ma}}$ and ${\rm{Kn}}$ on flow characteristics has been elucidated graphically in various sections. The flow properties showed significant variation with ${\rm{Ma}}$ and ${\rm{Kn}}$ and showed an increasing trend due to compressibility and viscous heating effects. The surface characteristics were observed to diminish as ${\rm{Ma}}$ increases, while they showed complex trends for various ${\rm{Kn}}$ . In all flow regimes, there was an appearance of flow recirculation. When chemical reactions were taken into consideration and compared with non-reacting flows, flow temperature was primarily influenced (which decreased due to energy absorption) compared to other properties. The results obtained are a complex interplay of the viscous heating, compression and rarefaction effects.

Particulate deposits in aero-engine turbines change the profile of blades, increase the blade surface roughness and block internal cooling channels and film cooling holes, which generally leads to the degradation of aerodynamic and cooling performance. To reveal particle deposition effects in the turbine, unsteady simulations were performed by investigating the migration patterns and deposition characteristics of the particle contaminant in a one-stage, high-pressure turbine of an aero-engine. Two typical operating conditions of the aero-engine, i.e. high-temperature take-off and economic cruise, were discussed, and the effects of particle size on the migration and deposition of fly-ash particles were demonstrated. A critical velocity model was applied to predict particle deposition. Comparisons between the stator and rotor were made by presenting the concentration and trajectory of the particles and the resulting deposition patterns on the aerofoil surfaces. Results show that the migration and deposition of the particles in the stator passage is dominated by the flow characteristics of fluid and the property of particles. In the subsequential rotor passage, in addition to these factors, particles are also affected by the stator–rotor interaction and the interference between rotors. With higher inlet temperature and larger diameter of the particle, the quantity of deposits increases and the deposition is distributed mainly on the Pressure Side (PS) and the Leading Edge (LE) of the aerofoil.

This paper consists of CFD and experimental study for shell projectile at angle of attack 00 and at various Mach numbers (0.7, 0.9, 1.0, 1.5 and 2.0), without spin effect. Passive method of base modification is used to reduce base drag. The goal of this study is to reduce base drag by utilising a base bleed approach called the Inward Turning Base Bleed Method (IWTB). The concept of IWTB is to draw relatively high-pressure air behind the driving band and allow it to pass through the bleed holes and direct it into the low-pressure base. This will in turn raise the base pressure and lower the base drag. The study comprises of (i) basic model with a boattail angle of $8^{\circ}$ , (ii) eight different cavity models having different lip thickness and depth thickness ratios and (iii) nine different IWTB configurations for the optimised cavity model. At first, cavity model is optimised by varying the lip thickness and depth thickness ratio. Later IWTB parameters such as the bleed hole entry angles $10^{\circ}$ , $12^{\circ}$ & $15^{\circ}$ , exit bleed hole angles $30^{\circ}$ , $45^{\circ}$ & $60^{\circ}$ , bleed hole diameters 3mm, 4mm, 5mm & 6mm and number of bleed holes i.e., 4 & 8 were studied and optimised. Based on this study, the model with 8 holes, 3mm base bleed hole diameter, $15^{\circ}$ entry and $60^{\circ}$ exit angle gives comparatively lower base drag. CFD result shows, in supersonic region the base drag reduction for optimised shell projectile is $3.08\%$ ; at Mach number 0.9 the base reduction is $75.63\%$ ; and in subsonic region it reduces to $23.53\%$ . CFD results were compared with experimental result, is found to be good and the differences lies within $3.85\%$ .

To obtain the optimal solution for the performance of the turbofan engine using infrared stealth technology, an engine mathematical model with a backward infrared radiation intensity calculation module was established. The effects of infrared suppression measures on the performance of turbofan engines were analysed. Based on the multi-objective particle swarm optimisation (MOPSO) algorithm, the optimal solution for the performance in the cruise state of the reference engine refitted with the infrared radiation suppression module was obtained; Further, through the multiple design points (MDPs) concept, the thermal cycle optimisation design of the turbofan engine was carried out. The results show that the integrated fully shielded guiding strut (IFSGS) with air film cooling had the ideal infrared suppression effect. Compared with the reference engine refitted with infrared radiation suppression module, the engine after cycle optimisation design could obtain better infrared stealth performance.

The ground delay program (GDP) is a commonly used tool in air traffic management. Developing a departure flight delay prediction model based on GDP can aid airlines and control authorities in better flight planning and adjusting air traffic control strategies. A model that combines the improved sparrow search algorithm (ISSA) and Multilayer Perceptron (MLP) has been proposed to minimise prediction errors. The ISSA uses tent chaotic mapping, dynamic adaptive weights, and Levy flight strategy to enhance the algorithm’s accuracy for the sparrow search algorithm (SSA). The MLP model’s hyperparameters are optimised using the ISSA to improve the model’s prediction accuracy and generalisation performance. Experiments were performed using actual GDP-generated departure flight delay data and compared with other machine learning techniques and optimisation algorithms. The results of the experiments show that the mean absolute error (MAE) and root mean square error (RMSE) of the ISSA-MLP model are 16.8 and 24.2, respectively. These values are 5.61%, 6.3% and 1.8% higher in MAE and 4.4%, 5.1% and 2.5% higher in RMSE compared to SSA, particle swarm optimisation (PSO) and grey wolf optimisation (GWO). The ISSA-MLP model has been verified to have good predictive and practical value.

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.

This paper aims to present a vertical take-off and landing unmanned aerial vehicle (VTOL UAV) configuration and numerically simulate its flight transition from hover to cruise and from cruise to hover. It can tilt the canard and wing along with two attached propellers. Additionally, two fixed front propellers are pointing upwards. Multi-body equations of motion are derived for this concept of aircraft, which are used to compute the flight transition trajectory from hover to cruise configuration. Furthermore, a transition control algorithm based on gain scheduling is described, which stabilises the aircraft while it accelerates from hover to cruise, gradually tilting the wing along with its propellers, sequentially switching between equilibrium states, as the stability cost functions thresholds are reached. The transition control algorithm of the conceptual aircraft model is numerically simulated.

An integrated parametric model involving the design of propulsion system, airframe and flight mission is presented. Based hereon, the carbon dioxide (CO2) emission characteristics of advanced direct-drive turbofan and open rotor powered aircraft are analysed against pertinent aircraft and propulsion system design parameters. In addition, initial concept-specific trend statements on nitrogen oxides (NOx) as well as propulsor noise emission characteristics are derived. The obtained results contribute to a better understanding of more appropriate aircraft design attributes for advanced system architectures.

This paper has proposed a new robust hybrid nonlinear guidance law, which accounts for a missile’s terminal line-of-sight (LOS) angle constraint, in order to intercept a non-cooperative maneuvering target. The proposed hybrid nonlinear guidance strategy consists of two phases; in the first phase, a guidance law named PIGL is derived from prescribed performance control and the inertial delay control method. In PIGL, a revised prescribed performance function is put forward, and a prescribed performance controller with unknown uncertainties is then derived. The controller smoothly drives both the LOS angle and its rate to a predesigned small region under unknown uncertainties that are induced by target’s maneuvers within a fixed time. Then, a guidance law named SIGL is activated, which is derived from sliding mode control and inertial delay control. By driving the desired sliding mode variable to zero within a finite time, the SIGL guidance law is able to achieve high terminal interception accuracy. The robustness of both of the proposed sub-guidance laws has been proved explicitly in this paper. The hybrid guidance law has the advantage of a tunable convergence rate of the LOS angle and the rate of the LOS angle at the beginning period, by which an excessive large initial maneuver can be avoided. Meanwhile, the hybrid guidance law also has the advantage of lower sensitivity to errors in the estimation of the time-to-go.

The 150th anniversary of the Royal Aeronautical Society offers an ideal opportunity to reflect on the spirit of innovation and collaboration fostered by the Society and its members over the past century and a half. Dr Jean Botti, an engineer with 31 patents to his name and Chief Technical Officer at Airbus Group, reflects on key innovation milestones for both Airbus Group and the industry as a whole. He also discusses the benefits of collaboration between the RAeS members and industry, and looks forward to an exciting new era of discovery – from electric flight (e-flight) to ‘smarter skies’ and the future development of new modes of flight which can only be imagined today.

A simple method is presented for calculating the heat transfer rate through a compressible laminar boundary layer. The temperature of the wall is assumed to be uniform, the viscosity being proportional to the absolute temperature with a ratio which may vary in an appropriate manner with position along the wall; the Prandtl number is arbitrary but greater than about 0·5.

The method assumes, following Lighthill, that heat transfer rates are determined mainly by the form of the velocity field in regions close to the wall. After applying the Ulingworth-Stewartson transformation the velocity is expanded in powers of the distance Y normal to the wall, three terms being retained. The first term, proportional to the skin friction, is zero near to a position of boundary layer separation. The second term is proportional to the pressure gradient and also to the wall temperature. Accordingly it can become close to zero when the wall is sufficiently cooled. The third term becomes important only when the first two are simultaneously close to zero and is proportional to the heat transfer rate. Since the three terms cannot, in fact, ever be all zero at the same position they form a uniformly valid non-trivial approximation to the velocity close to the wall.

With this velocity profile, and following similarity arguments first given by Liepmann, it is possible to reduce the integrated thermal energy equation to a first-order ordinary differential equation for the heat transfer rate, which is easily solved.

A comparison with an accurate solution of Poots, for a particular pressure gradient and wall temperature, yields agreement to within 1 per cent at one third of the distance from leading edge to separation and 3 per cent at twice this distance, with the error rising to 16 per cent at separation where the heat transfer rate is, of course, very low. The modesty of this error is very encouraging.

The result derived in a previous paper for the primary effect of the jet mixing on the thrust is confirmed, both by an investigation using momentum principles and by a quantitative examination of the sink effect. Boundary layer behaviour still cannot be predicted with accuracy, but it is suggested that, even under cruising conditions, a jet flapped aerofoil may exhibit unconventional properties.

A numerical method which solves the time-dependent Navier-Stokes equations for plane, two-dimensional, compressible flow problems is presented. The method was applied to obtain the steady-state flow field about a flat-based 20° wedge at a free-stream Mach number of 6.05, a free-stream Reynolds number (based on the base height of the wedge) of 1.41 × 104, and a Prandtl number of 0.75. Two wedge surface temperature boundary conditions were investigated: adiabatic and isothermal walls. Convergence of the numerical solutions was demonstrated using two different finite-difference meshes. Numerical errors in the static pressure, velocity, and static temperature were generally less than 8 per cent in the wake of the body. Comparisons with experimental results and other independent theories were also made. For the isothermal wedge, calculated pressures on the surface of the body (including its base), static pressure profiles, pitot pressures profiles, and velocity profiles in the near wake were in agreement with corresponding experimental data. A comparison of static temperature profiles in the near wake of the isothermal wedge produced significant differences between numerical and experimental results near the wake shock. It was found that these differences could be explained in terms of uncertainties in the accuracy of the experimental measurements and thermal boundary conditions on the wedge surface. For the adiabatic wedge, the ramp boundary layer profiles and pressure distribution were in agreement with theoretical predictions coming from hypersonic interaction theory.

The theory of aerofoils in unsteady flow, which has made substantial progress in the last decade due largely to the ground work of Küssner and his co-workers, is presented here in a form suitable for application in aeroelastic problems, particularly those concerned with the dynamic loads on aircraft arising from gusts.

Exact expressions are given for the aerodynamic lift and moment for an oscillating aerofoil as well as for the case of arbitrary motion through disturbed air. The expressions for the latter case involve two special functions, generally referred to as Wagner and Küssner functions. Exact values of these functions are tabulated together with useful approximations. The problem of a wing-tail combination is discussed and a method of solution indicated. The bibliography at the end of the paper lists the most important publications in this field.

This article reports on the development of a fast method for generating the aerodynamic database for subsonic flow over a missile. At present, this is typically achieved using RANS-based CFD, which is expensive for the complex missile geometries and the multiplicity of operating conditions to be evaluated. The presented reduced-order model (ROM) provides a reasonably accurate prediction of the aerodynamic coefficients of the missile (and, in fact, the full flow field around it) within half a minute. In particular, in the interpolative regime, prediction errors for all coefficients are typically less than 1% of their respective maximum values encountered in the database, with extrapolation incurring more error. The empirical approach ‘learns’ from the CFD solutions calculated for a few operating conditions, and then is able to make predictions for any other condition within a feasible set. The learning employs proper orthogonal decomposition (POD) to characterise the most important features of the flow. The prediction is posed as an optimisation problem that aims to find the flow solution as a linear combination of the above POD modes that minimises the residual of the governing equations. Innovations on the prevailing POD-ROM approach include novel implementation of boundary conditions, simplified computation of the aerodynamic coefficients, and a procedure for choosing modelling parameters based on extensive cross-validation. Challenges overcome in application to a problem of industrial relevance are discussed.

Our previous speaker, Mr. Blyth, in his paper on the Design of Incentive Contracts, has given us a lucid explanation of the basic principles of incentives and examples of cost, performance and delivery conditions which could be used as incentives and the objectives to which they could be directed.

Following on this paper, one could initiate a discussion on the basic philosophy behind the use of incentives, whether the incentives as stated are likely to be achieved, and propose various alternative methods of achieving these objectives without using contractual conditions. I have no doubt that many points such as these will be raised later during the discussion.

The correct design of intermediate vertical stiffeners on web plates subjected to shear becomes very important when the web plates are designed to operate at loads close to their buckling loads. This paper presents details of an extensive series of tests conducted on stiffened web plates subjected to shear. From the analysis of the results obtained from these tests, new empirical relationships between the flexural rigidity and spacing of the intermediate stiffeners and the buckling stress of the stiffened web plate have been obtained.

One interesting and important feature of these new relationships is that they define more clearly than hitherto the difference in the behaviour of single-and double-sided stiffeners.