In an effort to reduce the negative impact of civil aviation on the human environment, trajectory optimisation techniques have been used to minimise the single event impact of noise and gaseous emissions of departures on communities in the vicinity of airports. For this purpose, the earlier developed trajectory optimisation tool NOISHHH has been adapted to design departure trajectories optimised for environmental criteria, based on area navigation. The new version of NOISHHH combines a noise model, an emissions inventory model, a geographic information system and a dynamic trajectory optimisation algorithm to generate flight paths with minimised environmental impact. Operational constraints have been introduced to ensure that the resulting flight paths are fully compliant with the guidelines and regulations that apply to the design of standard instrument departures and the use of area navigation. To illustrate the capabilities of the new version of NOISHHH, two numerical examples are presented, which are both redesigns of standard instrument departures currently in use at Amsterdam Airport Schiphol.

An analysis is made of the general two-cell thin-walled cylindrical tube subjected to longitudinal tension, bending and torsion. The walls of the tube may be of fibre composite with an asymmetric lay-up resulting in a coupling between direct and shear stresses and strains. The two-cell tube differs from the single-cell tube in that the tension axis and the principal axes of bending are not necessarily aligned to the c.g. and the principal moments of inertia of the longitudinal stress bearing material.

Recently, transformable Unmanned Aerial Vehicles (UAVs) have become a subject of great interest in the field of flying systems, due to their maneuverability, agility and morphological capacities. They can be used for specific missions and in more congested spaces. Moreover, this novel class of UAVs is considered as a viable solution for providing flying robots with specific and versatile functionalities. In this paper, we propose (i) a new design of a transformable quadrotor with (ii) generic modeling and (iii) adaptive control strategy. The proposed UAV is able to change its flight configuration by rotating its four arms independently around a central body, thanks to its adaptive geometry. To simplify and lighten the prototype, a simple mechanism with a light mechanical structure is proposed. Since the Center of Gravity (CoG) of the UAV moves according to the desired morphology of the system, a variation of the inertia and the allocation matrix occurs instantly. These dynamics parameters play an important role in the system control and its stability, representing a key difference compared with the classic quadrotor. Thus, a new generic model is developed, taking into account all these variations together with aerodynamic effects. To validate this model and ensure the stability of the designed UAV, an adaptive backstepping control strategy based on the change in the flight configuration is applied. MATLAB simulations are provided to evaluate and illustrate the performance and efficiency of the proposed controller. Finally, some experimental tests are presented.

The class of aircraft known as gyroplanes (or autogyros) helped to pave the way for the development of the helicopter. However, they have found no application in contemporary commercial or military aviation. It is in recreational or sport flying that the gyroplane has proved popular. Most if not all designs are however homebuilts, and as a consequence little analysis of any significance has been conducted on the flight mechanics of these aircraft. This Paper presents an analysis of the longitudinal stability of gyroplanes. Simple consideration of the basic forces and moments that might influence stability, in the context of elementary rotor behaviour, is used to assess the fundamental nature of gyroplane longitudinal stability. This is quantified by the application of a sophisticated generic rotorcraft mathematical model to the gyroplane problem. It is concluded that the basic configuration can be longitudinally stable with appropriate design, and that the rotorspeed degree of freedom must also be included when modelling the aircraft.

In many cases of the flow through porous screens, one may consider it to be made up of a number of jets passing through the openings of the screen. These jets are separated by a series of wakes behind the solid parts of the screen. The majority of investigations on the flow through such screens have been concerned with the measurement of pressure drop and its variation with different flow conditions; it has been assumed that the pressure is discontinuous at the screen itself and that the pressure drop coefficient Δp/½ρυ2 provides sufficient information, where Δp is the pressure drop across the screen, ρ the density of the fluid, and υ the velocity of approach to the screen.

A method for the numerical solution of the two-dimensional Euler equations on unstructured grids has been developed. The cell-centred symmetric finite-volume spatial discretisation is applied in a general formulation that allows the use of arbitrary polygonal computational cells. The integration in time, to a steady-state solution, is performed using an explicit, multi-stage procedure, with standard convergence acceleration techniques such as local time stepping, enthalpy damping and implicit residual smoothing. Accuracy of solution, in terms of minimising spurious entropy production, is achieved through careful treatment of the artificial dissipative terms near boundaries. Standard test cases for both subcritical and supercritical flows, including single- and multi-element aerofoils have been used to validate the method.

Ultrasonic guided wave techniques have great potential for structural health monitoring applications. Appropriate mode and frequency selection is the basis for achieving optimised damage monitoring performance. In this paper, several important guided wave mode attributes are introduced in addition to the commonly used phase velocity and group velocity dispersion curves while using the general corrosion problem as an example. We first derive a simple and generic wave excitability function based on the theory of normal mode expansion and the reciprocity theorem. A sensitivity dispersion curve is formulated based on the group velocity dispersion curve. Both excitability and sensitivity dispersion curves are verified with finite element simulations. Finally, a goodness dispersion curve concept is introduced to evaluate the tradeoffs between multiple mode selection objectives based on the wave velocity, excitability and sensitivity.

Aeroelastic analysis is a critical area of the aircraft design process, as a good understanding of the dynamic behaviour of the wing structure is essential to safe operation of the vehicle. The inevitable inaccuracies present in the modelling of such phenomena impose mass penalties, as large safety margins are necessitated, which in turn lead to overly stiff designs. In an effort to reduce the uncertainty in analysis methods, fully coupled CFD and structural models are under widespread development. This paper describes the results produced by such a system for a series of test cases based on the AGARD445.6 and MDO wings. Results relating to the latter are of particular interest, as significant variations were found to be produced by the different methodologies used in previous studies, the precise cause of which could not be isolated. In an effort to provide this isolation, a detailed description of the method used is given, including the interpolation scheme between the structural model and the aerodynamic surface, and particular attention is given to the issue of aerofoil shape preservation.

Based on two-phase flow theory, an Eulerian method to simulate rime ice accretions on an aerofoil has been developed. The SIMPLE (semi-implicit method for pressure linked equations) algorithm on a collocated grid is employed to solve the governing equations for the airflow. In order to simulate droplets impinging on an aerofoil, a permeable wall is proposed to solve the governing equations for supercooled droplets. The collection efficiency and impingement limits are obtained from the droplets’ flowfield. The process of ice accretion is simulated using the assumption that ice accumulates layer-by-layer and the ice shape is predicted with the assumption that ice grows in the direction normal to the aerofoil surface. The rime ice accretions on a NACA0012 aerofoil at 0° and 4° angles-of-attack have been investigated and there is agreement between the simulated results and previously published experimental data. The change of the pressure coefficient along the iced aerofoil is also analysed.

The cylindrical tankage which forms part of the primary structure of missiles and space vehicles may be subjected to high compressive axial and bending loads. Since allowable compressive stresses for monocoque construction are generally low because of small bending rigidity and sensitivity to geometric imperfection, it is advantageous to use some type of wall stiffening. One such choice is the integrally milled-out 45° waffle. This consists of a square gridwork of ribs which form 45° helices over the cylindrical surface. The panel is generally milled out in the flat and then bent up to form a circular cylinder. An advantage of the 45° construction is the exclusion of buckling modes between hoop reinforcement; such modes are frequently critical for the 0° to 90° type of stiffening.

This paper reviews findings on the type, morphology and constitutive behaviour of impact damage zones during loading after impact and their effect on the laminate strength and stability. The paper is limited to tape prepreg based monolithic laminates, although some similarities exist with impact damage in textile based laminates. Damage zones have a complex geometry with several damage types, which results in an interaction and competition between different failure mechanisms, e.g. local and global buckling, compressive failure, and delamination growth. Hence, simplified damage models may provide incorrect predictions of the failure load and failure mechanisms after impact. The constitutive behaviour of damage zones has been studied experimentally in tension and compression using an inverse method, and the results have been compared with detailed FE models of a generic impact damage. The paper is concluded with a discussion on analytical and computational models to predict the resulting strength of impacted laminates.

Inlet unstart boundary is one of the most important issues of the hypersonic inlet and is also the foundation of the protection control of a scramjet. To solve this problem, the 2D internal steady flow of a 2D mixed internal/external compression hypersonic inlet was numerically simulated at different freestream conditions and backpressures with a RANS (Reynolds-Averaged Navier-Stokes) solver using a RNG (Renormalisation Group) k-ε turbulence model, and two different inlet unstart phenomena were analysed. The dimensional analysis method was introduced to find the essence variables describing the inlet unstart boundary based on “numerical experimental” data in this paper. The dimensionless pressure ratios of the forebody and isolator were analysed respectively. The results show that the unstart boundary of the 2D mixed hypersonic inlet is determined by M0, α and Re0. Pressure ratio π increases with M0 increasing, and it increases firstly and decreases then with α increasing. Pressure ratio π increases with Re0 increasing. Re0 (Re0 < 2 × 107) has a major effect on π and Re0 (Re0 > 2×107) has little effect on π.

When the stagnation temperature of a perfect gas increases, the specific heats and their ratio do not remain constant and start to vary with the temperature. The gas remains perfect; its state equations remain valid, so it can be named as calorifically imperfect gas. The aim of this research is to develop the necessary thermodynamic and geometrical equations and to study the supersonic flow at high temperature, lower than the dissociation threshold. The results are found by the resolution of nonlinear algebraic equations and integration of complex analytical functions where the exact calculation is impossible. The dichotomy method is used to solve the nonlinear equations and Simpson’s algorithm for the numerical integration applied. A condensation of the nodes is used. The functions to be integrated have a high gradient at the extremity of the interval of integration. The comparison is made with the calorifically perfect gas to determine the error. The application is made for air in a supersonic nozzle.

There is a growing interest in the use of alternative fuels in gas turbine engines to reduce emissions. Testing of alternative fuels is expensive when done on a large-scale gas turbine engine. In this study, a re-commissioned small gas turbine auxiliary power unit (APU) has been used to test various blends of Jet A-1, synthetic paraffinic kerosene (SPK) and diesel with as well as eight other novel fuels. A detailed analysis of performance, gaseous emissions and particulate emissions has been presented in this study. It is observed that aromatic content in general as well as the particular chemical composition of the aromatic compound plays a vital role in particulate emissions generation. SPK fuel shows substantially lower particulate emissions with respect to Jet A. However, not all the species of aromatics negatively impact particulate emissions. Gaseous emissions measured are comparable for all the fuels tested in this study.

This paper presents a dynamic analysis of a single-line kite with two degrees of freedom. A Lagrangian formulation is used to write convenient equations of motion. The equilibrium states of the system and their stability are studied; Eigenvalues and eigenmodes are calculated by using linear theory. The stability in the parametric plane δ – W0 is discussed, where δ defines the bridle geometry and W0 is wind velocity. The system goes through a Hopf bifurcation and periodic branches of solutions appear. The orbits and their stability have been calculated numerically using Floquet theory and wind velocity seems to play an important role in their existence. Finally the kite response against gusts is considered and an open loop control system developed to keep the flight altitude invariant under changing atmospheric conditions. Modifying the bridle’s geometry seems to be a convenient way to control a kite’s performance.

Experiments are performed on the effects of a turbulent boundary layer on surface pressure and heat transfer, using a cylindrical model fitted with a variety of flares at Mach 5. The flow over compression corners of up to 20° (below the expected incipient separation value) is examined. Liquid crystal thermography indicates an unexpected peak in heat flux close to the corner. It is suggested that this is due to a small scale separation occurring earlier, i.e. at smaller corner angles, than previously observed. An estimate of the size of such a separation bubble is made and it is suggested that this is a predominantly laminar phenomenon.

Effects of boundary conditions on the performance of compact oscillatory momentum and vorticity generators, commonly known as ‘synthetic jet’, ejecting a train of vortex-pairs into still air, were studied experimentally. The different boundary conditions altered the near-device entrainment process of the zero-net-mass-flux actuator. The measurements included hot-wire and Particle Image Velocimetry, cavity pressures and temperatures.

When the actuator operates in still air, a quasi-2D vortex pair is generated due to the extreme shear at the edges of the fluid slug ejected during the blowing stage of each excitation cycle. The vorticity flux exiting the slot determines the resulting vortex-pair circulation. The threshold slot exit velocity, for the current configuration and operating conditions, determines if the vortices will be sucked into the actuator’s cavity or be released. Once released, the vortex convection speed approximately scales with the peak velocity at the slot exit. However, the normalised convection velocity increases with the slot Reynolds number.

When even a very short extension is attached to one ‘lip’ of the actuator exit, the jet is deflected in the direction opposite the extended lip, due to the restriction on the entrainment process. When long, one ‘lip’ extension is attached, such that the vortex pair is ejected parallel to a plate, the coherence of the vortices improve, their phase speed and magnitude decrease.

The effects of high-frequency excitation, ejected perpendicular to a wall into still air, were also investigated. It was found that the presence of the plate does not have a measurable effect on the wall normal excitation, indicating that the majority of the entrainment is taking place from the forward 180° of the actuator exit plane. When the slot is inclined to the surface at a shallow angle of 30 degrees, an unsteady wall jet is formed, transferring momentum along the wall. This is a direct result of the symmetry break, altering the relative magnitudes of the vortex pair.

The ground vortices generated by an intake under both headwind and crosswind configurations have been investigated using computational and experimental approaches. The flow field of a scale-model intake was experimentally studied using stereoscopic particle image velocimetry to measure the ground vortex in conjunction with induct total pressure measurements for the internal flow. The computational predictions were performed using an unsteady Reynolds averaged Navier-Stokes approach. The experimental results show that under crosswind conditions a single ground vortex forms which becomes stronger as the crossflow velocity is increased. Under headwind conditions the measured ground vortex strength initially increases with freestream velocity before it reaches a local maximum and then reduces thereafter. The computations also exhibit the same characteristics and show good agreement with the measurements for some configurations. Based on the predictions, the complex flow field topology is investigated and a detailed flow model of the vortex flow field under crosswind conditions is proposed.

This paper considers the dynamic stall vortex of importance in helicopter rotor aerodynamics and discusses previous measurements of its convection speed. It emerges that an anomaly exists between the available data sets, i.e. that some workers find that the convection speed is dependent upon the aerofoil motion, while others find that this is not the case. Measurements of the convection speed from data gathered at Glasgow University for a variety of aerofoil shapes and motion types are then presented, which support the conclusion that the dynamic stall vortex convection speed is independent of aerofoil type and motion type to a first order.

The degree of complexity and randomness of aircraft tracking which can be safely managed is a fundamental question for a given airspace design. To answer this question and to provide optimum conflict resolution strategies, the nature of conflict needs to be thoroughly understood. This paper reviews traditional conflict models showing how these models may be integrated into a more unified model based on computational geometry. This mathematical approach is being investigated for the specification and management of future airspace, achieving a more solid architectural basis for airspace design. This approach has the potential to significantly reduce the subjectivity when attempting to model controller and pilot workload, an important requirement in future designs.