The ever-changing battlefield environment, as well as the emergence of global command and control architectures currently used by armed forces around the globe, requires the use of robust and adaptive technologies integrated into a reliable platform. Unmanned Combat Aerial Vehicles (UCAVs) aim to integrate such advanced technologies while also increasing the tactical capabilities of combat aircraft. This paper provides a summary of the technical and operational design challenges specific to UCAVs, focusing on high-performance, and stealth designs. After a brief historical overview, the main technology demonstrator programmes currently under development are presented. The key technologies affecting UCAV design are identified and discussed. Finally, this paper briefly presents the main issues related to airworthiness, navigation, and ethical concerns behind UAV/UCAV operations.

Donnell type equilibrium and stability equations are derived for stiffened thin conical shells. The stiffeners are considered closely spaced and are therefore assumed to be “distributed” over the whole surface of the shell. In the proposed theory the stiffeners and their spacing may vary in any prescribed manner, but here only equal and equally spaced stiffeners are dealt with. The force- and moment-strain relations of the combined stiffener-sheet cross section are determined by the assumption of identical normal strains at the contact surface of stiffener and sheet.

The stability equations are solved for general instability under hydrostatic pressure by the method of virtual displacements. The solution used earlier for unstiffened conical shells, which satisfies some of the boundary conditions of simple supports only approximately, is again applied here. The effect of this incomplete compliance with boundary conditions is shown to be negligible by consideration of “boundary work”. The solution proposed for stiffened conical shells involves the concepts of “correcting coefficients” and minimisation of corresponding “error loads”.

Typical examples are analysed and the effect of eccentricity of stiffeners is investigated. Simplified approximate formulae for the critical pressure of frame-stiffened conical shells are also proposed.

This paper presents an overview of an advanced, conceptual wing-box weight estimation and sizing model for transport aircraft. The model is based on linear thin-walled beam theory, where the wing-box is modelled as a simple, swept tapered multi-element beam. It consists of three coupled modules, namely sizing, aeroelastic analysis, and weight prediction. The sizing module performs generic wing-box sizing using a multi-element strategy. Three design cases are considered for each wing-box element. The aeroelastic analysis module accounts for static aeroelastic requirements and estimates their impact on the wing-box sizing. The weight prediction module estimates the wing-box weight based on the sizing process, including static aeroelastic requirements. The breakdown of the models into modules increases its flexibility for future enhancements to cover complex wing geometries and advanced aerospace materials. The model has been validated using five different transport aircraft. It has shown to be sufficiently robust, yielding an error bandwidth of ±3%, an average error estimate of -0·2%, and a standard error estimate of 1·5%.

A method is presented for determining the swirling compressible flow through a nozzle, given conditions at a reference section. The principal assumption is that changes in the nozzle cross-sectional area are sufficiently gradual for the radial velocity component to be neglected at each section, i e, the usual assumption of one-dimensional compressible flow theory. This method is used to determine choked-flow conditions in the case where there is solid-body rotation at the throat for a range of swirl intensities with the ratio of the specific heats taking various values. Mass-flux coefficients and impulse functions are determined. Sonic surfaces, velocity profiles and other characteristics of interest are also presented. An approximate analysis valid for low swirl intensities is developed and analytical formulae are derived for most quantities of interest. The main conclusion of practical importance is that the introduction of swirl to compressible nozzle flows need not lead to a significant reduction in specific thrust. Further exploration of the possible effects of swirl on noise during the relatively short take-off and landing periods cannot therefore be ruled out on the grounds that swirl would lead to excessive thrust losses.

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.

The obtuse corner of the skew plate is a singular point with a stress singularity growing in intensity with an increase in the angle of skew. For regaining the full rate of convergence of the finite element method the mesh around the singular point should be properly refined or the leading singular terms included in the finite element scheme. In any event a large number of unknowns may be required. For obtaining an accuracy of 1% in the 10th Eigenvalue of a clamped square plate discretised with the 16-degrees-of-freedom elements used in the present calculations, about 300 variables are required. For computational convenience calculations were carried out here with a uniform mesh requiring even a larger number of unknowns.

It is shown from elementary considerations that the lift coefficient of a thin two-dimensional wing at zero incidence, with a narrow high-velocity jet of momentum coefficient Cj issuing from its trailing edge at a (small) downward inclination τ, is given by

and the loading on the chord line (0<x<1) by

(except in a certain neighbourhood of the trailing edge), for small values of Cj. These formulae agree well with known measurements. Interpolation formulae for derivatives of CL at larger values of CJ based on earlier work, are also given:

Three computer programmes have been written for the stability analysis of both isotropic and orthotropic, multi-plate structures in compression. The bases of these programmes are the exact and approximate Finite Strip methods. A considerable number of results for a range of multi-plate structures fabricated from both carbon fibre reinforced (orthotropic) and aluminium (isotropic) materials are presented in graphical form. These graphs serve to illustrate not only the influence of material orthotropy on the buckling response, but also the possible savings in weight that may be achieved by employing orthotropic materials in place of conventional isotropic materials.

In the flow over high-lift aerofoils there is a strong interaction between the wake from one element and the boundary layer on another element immediately downstream. A number of experimental investigation of these types of flow have been made but these studies do not allow the effects associated with the various parameters of the aerofoil to be seen easily. This paper presents the results of extensive measurements of mean and fluctuating quantities made in merging layers in zero and adverse pressure gradients downstream of aerofoils of various shapes. A general description of the emerging process is given and the effects of wake characteristics and pressure gradient are considered. The paper also contains a discussion of the modelling of the Reynolds shear stress in the merging layer.

A method to predict the effects of rotor icing on the flight characteristics of a UH-60A helicopter is presented. By considering both natural ice shedding and different types of ice accretion due to local temperature variations on the blade surface, an improved rotor icing model was developed. Next, the effects of icing on rotor force, torque and flapping were incorporated in a nonlinear helicopter dynamic model. Based upon icing design envelopes in cumuliform clouds, trim and stability characteristics were studied. Further development of the helicopter state-space model allowed control and handling qualities characteristics to be investigated with variation of the three icing-related cloud variables (atmospheric temperature, liquid water content, and median volumetric diameter). Results indicated that this method of evaluating rotorcraft icing is both feasible and useful.

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 aerodynamic shape optimisation of a micro air vehicle (MAV) wing is performed to obtain the basic wing geometrical characteristics which produce the maximum range and endurance requirements. Multhopp’s method based on Prandtl’s classical lifting line theory is used for the determination of the spanwise load distribution required during the optimisation process. The obtained lift and drag characteristics are used for the derivation of the range and endurance equations of an electrically driven micro air vehicle. The optimisation process is based on the modified feasible directions gradient based optimisation algorithm. Results are validated using wind tunnel measurements showing very good agreement. Results are also compared with solutions to the Navier-Stokes equations obtained with ANSYS-CFX finite elements using different turbulence models. These include the k-ε and the shear stress transport (SST) models as well as the Reynolds stress model.

One of the key challenges of the gas turbine community is to empower the condition based maintenance with simulation, diagnostic and prognostic tools which improve the reliability and availability of the engines. Within this context, the inverse adaptive modelling methods have generated much attention for their capability to tune engine models for matching experimental test data and/or simulation data. In this study, an integrated performance adaptation system for estimating the steady-state off-design performance of gas turbines is presented. In the system, a novel method for compressor map generation and a genetic algorithm-based method for engine off-design performance adaptation are introduced. The methods are integrated into PYTHIA gas turbine simulation software, developed at Cranfield University and tested with experimental data of an aero derivative gas turbine. The results demonstrate the promising capabilities of the proposed system for accurate prediction of the gas turbine performance. This is achieved by matching simultaneously a set of multiple off-design operating points. It is proven that the proposed methods and the system have the capability to progressively update and refine gas turbine performance models with improved accuracy, which is crucial for model-based gas path diagnostics and prognostics.

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.

Nishihara’s combined bending and torsion out-of-phase fatigue limit data are analysed. The Tresca shear stress failure criterion predicts strengths up to 30 per cent higher than observed. It thus appears that renewed attention should be given to the basic problem of developing reliable combined stress failure criteria. It is suggested that new test methods will be required for this purpose.

Two-dimensional photoelastic tests have been carried out on uni-axially loaded flat-plate specimens with two collinear edge slits, to investigate the effect of finite plate width on the elastic stress distribution. It was found that the effect of slitlength/ plate-width ratio on the elastic stress concentration at the end of the edge slit of length l was virtually the same as that for a central slit of length 2l in a plate of the same width, and could be adequately expressed by existing theories.

The long-term performance of aeronautical structures is typically discussed in terms of concepts such as structural integrity, durability, damage tolerance, fracture toughness, etc. These familiar concepts are usually addressed by considering balance equations, crack growth relationships, and constitutive equations with constant material properties, and constant or cyclically applied load conditions. Loading histories are represented by changing stress (or strain) states, only. But for many situations, especially associated with high performance aircraft, the local state of the material may also change during service, so that the properties used in those equations are functions of time and history of applied conditions. For example, local values of stiffness, strength, and conductivity are altered by material degradation to create ‘property fields’ that replace the global constants, and introduce time and history into the governing equations. The present paper will examine a small set of such problems and offer a construct for using related solutions to estimate future performance based on history of use and current material state, a concept typically called prognosis.

“Urban air vehicles” have been hailed as the next revolution in aviation. Prototypes of various sizes have been flown to demonstrate basic flight (hover and climb), but in most cases there is no demonstration of full flight capability, for example conversion from vertical to level flight (conversion corridor). There are proposals for vehicles in a wide range of scales: from drones specifically designed to deliver goods, to full size vehicles for manned transportation. Most of the concepts proposed include full electric propulsion, multiple (often convertible) rotors (ducted or un-ducted, counter-rotating), and widespread use of composite materials. Start-up companies are seeking funding with high-profile demonstrations in front of the media, but many unresolved technical problems are not been solved. Large aerospace companies have joined the fray. These initiatives are fuelling expectations that achieving the next milestone is within easy reach. This paper aims to fill some gaps in understanding and curb optimism. It takes a holistic view in order to establish a scientific basis for design, manufacturing, operations.

A large deflection theory is established for the bending and stretching of aeolotropic plates with elastic moduli that vary in the x, y, z directions. The theory is formulated in terms of two simultaneous non-linear differential equations, with variable coefficients, for the transverse deflection w and an Airy stress function F. Small deflection heterogeneous plate theory and some other sub-classes of plate problems are formulated by specialising the general non-linear plate equations. Effects of initial irregularities and of a variable elastic foundation are considered.

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.