A thermodynamically-based work potential theory for modelling progressive damage and failure in fibre-reinforced laminates is presented. The current, multiple-internal state variable (ISV) formulation, enhanced Schapery theory, utilises separate ISVs for modelling the effects of damage and failure. Damage is considered to be the effect of any structural changes in a material that manifest as pre-peak non-linearity in the stress versus strain response. Conversely, failure is taken to be the effect of the evolution of any mechanisms that results in post-peak strain softening. It is assumed, matrix microdamage is the dominant damage mechanism in continuous, fibre-reinforced, polymer matrix laminates, and its evolution is captured with a single ISV. Three additional ISVs are introduced to account for failure due to mode I transverse cracking, mode II transverse cracking, and mode I axial failure. Using the stationarity of the total work potential with respect to each ISV, a set of thermodynamically consistent evolution equations for the ISVs is derived. Typically, failure evolution (i.e. post-peak strain softening) results in pathologically mesh dependent solutions within a finite element method numerical setting. Therefore, consistent characteristic element lengths are introduced into the formulation of of the three failure potentials. The theory is implemented into commercial FEM software. The model is verified against experimental results from a laminated, quasi-isotropic, T800/3900-2 panel containing a central notch. Global load versus displacement, global load versus local strain gauge data, and macroscopic failure paths obtained from the models are compared to the experiments. Finally, a sensitivity study is performed on the failure parameters used in the model.

A diffusing S-duct aircraft intake system is computationally studied for the effects of inlet icing and wall heat transfer on engine face distortion. A Reynolds-Averaged Navier-Stokes (RANS) code with k -ω SST turbulence model is used to simulate the compressible viscous flow in the duct. The glaze ice accretion on the inlet lip is simulated as a fixed protrusion from NASA LEWICE3D code. The shape and size of the ice on the inlet lip are assumed to remain constant even in the case of the heated wall. The case of zero external heat transfer is modeled by the adiabatic wall boundary condition, and constant-wall temperature is used to simulate the heated wall effect. The freestream Mach number of 0·85 and glaze ice condition on the inlet lip produce massive flow separation at the lip and internal shock cell structure in the S-duct. The heated wall creates additional vorticity and thus increases the engine face distortion level. The area-averaged total pressure distortion at the engine face in a freestream Mach number of 0·85 is increased by ~7% in the heated wall case as compared to the adiabatic flow.

Analytical methodology is presented to conduct dynamical assembly of simple low order nonlinear responses for system synthesis and prediction using Volterra theory. The procedure is set forth generically and then applied to several atmospheric flight examples. A two-term truncated Volterra series, which is enough to capture the quadratic and bilinear nonlinearities, is developed for first and second order generalised nonlinear single degree of freedom systems. The resultant models are given in the form of first and second kernels. A parametric study of the influence of each linear and nonlinear term on kernel structures is investigated. A step input is then employed to quantify and qualify the nonlinear response characteristics. Uniaxial surge and pitch motions are presented as examples of the low order flight dynamic systems. These examples show the ability of the proposed analytical Volterra-based models to predict, understand, and analyse the nonlinear aircraft behaviour beyond that attainable by linear-based models. The proposed analytical Volterra-based model offers an efficient nonlinear preliminary design tool in qualifying the aircraft responses before computer simulation is available or invoked.

One of the most resent efforts in aircraft design is the replacement of aluminium structures by carbon fibre reinforced polymer composites. Due to lower material and manufacturing costs, doubly curved shapes covering big areas are preferred over simpler surfaces which integrate stiffening profiles. In this context, CAD parameterisation of surfaces allows design solutions by means of classical shape optimisation. Related geometrical parameters are manipulated towards optimal design, generating innovative geometries and detailing.

The presented structure is optimised by reducing the overall weight. The final optimum is guided using stability and strength restrictions in order to assure the safety of the component. Geometrical considerations are also included due to operational reasons. A hierarchical design procedure is developed which results in a work flow from preliminary ‘parameter-free’ form finding motivated by solving the minimal surface problem. The geometrical model for optimisation is recovered by generating B-Spline surface patches to preserve continuity requirements over large regions. The number of geometrical coefficients are defined by the accuracy in surface generation and the required freedom in surface control. The hierarchical approach reduces the possibilities of ending with an unsatisfactory optimum when several local minima characterise the non-linear problem, as it is usually the case in shape optimal design. A geometrical non-linear analysis is used to verify the performance of the optimum.

This article presents the findings of a study to investigate the impact perforation resistance of sandwich structures. The dynamic response of sandwich panels based on PVC foam cores has been evaluated by determining the energy to perforate the panels. The impact response of the sandwich structures was predicted using the finite element analysis package Abaqus/Explicit. The validated FE models were also used to investigate the effect of oblique loading and to study the impact response of sandwich panels subjected to a pressure differential equivalent to flying at an altitude of 10,000m.

Low velocity impact testing has shown that the energy to perforate the sandwich panels is dependent on the properties of the core. It has been shown that increasing the density of the crosslinked PVC foams by a factor of two yielded a 600% increase in the perforation resistance of the sandwich structures. At higher densities, the crosslinked foam sandwich structures offered a superior perforation resistance to the linear PVC structures. The numerical analysis accurately predicted the perforation energies of the sandwich panels, as well as the prevailing failure mechanisms following impact. Finally, it has been shown that sandwich panels impacted at high altitude offer a similar perforation resistance to those tested at sea level.

The flowfield around a supersonic projectile using a pin actuator control method has been predicted using computational fluid dynamics. It has been predicted using both viscous and inviscid methods for a number of positions. Both methods showed that an optimal longitudinal position exists. However, the inviscid model over-predicted the lateral acceleration due to the difference in shock formation around the pin between the two approaches. The optimal location was predicted independent of solver, however the higher-fidelity solver predicted lower achievable lateral accelerations. This is due to the viscous interactions caused by the pin. The effect of projectile orientation has shown that shielding the pin leads to reduced effectiveness due to the wake of the fin enveloping the pin. When the pin is exposed to onset flow, the forces achieved are increased. There is also an increase in the achievable forces and moments with increasing Mach number.

The satellite attitude determination approach based on the Extended Kalman Filter (EKF) has been widely used in many real applications. However, the accuracy of this method largely depends on the fitness of measurement model. We aim to analyse the influence of measurement errors to the accuracy of EKF based attitude determination approach in this paper. The measurement errors, which are divided into structural error and nonstructural error by their influences, are analysed in principle. In the setting of the combination of star sensors and gyros, according to the property of innovation, we employ the technique of correlation test to analyse the influences of different kinds of measurement errors. Experimental results demonstrate the effectiveness of our previous analysis.

In this paper, helicopter rotor blades are analysed in hover using computational fluid dynamics (CFD) coupled with a structural model. The method relies on a mesh deformation algorithm that allows for exchange of forces and deformations between a beam-based finite-element model and the fluid flow volume mesh. The method is demonstrated against experimental data, and the aerodynamic predictions appear to improve when the aeroelastic model is used. For all employed cases the flexibility of the method allows the CFD mesh deformation to be spread over the computational domain in a controlled fashion. The influence of the aeroelastic deformations on the blade loads was limited yet evident on the rotor performance. The lack of adequate test cases and experiments for validation of CFD/CSD methods is also highlighted.

Numerical simulations were performed to analyse the possibility of damping abrupt incoming free stream disturbances upon a porous aerofoil using an unsteady Reynolds-averaged Navier-Stokes (URANS) model. To mimic the turbulence disturbance levels that are typically encountered in the atmosphere, two flow configurations were considered. In the first configuration, the unsteadiness of the flow was created with vortices shed from a circular cylinder installed ahead of a WTEA-TE1 aerofoil. The continuous von Kármán shedding vortices contained within the cylinder wake were convected downstream and projected upon the aerofoil. In the second configuration, an instantaneous pair of discrete vortices was created by a rotational snapping of a flat plate, installed upstream of the aerofoil. Solid and porous aerofoil configurations, with porosity settings of 11 and 22%, were applied on 50% of the chord of the aerofoil starting from the leading edge. Both steady and unsteady flow simulations were performed to assess the performance of the porosity under steady and unsteady effects. The steady state flow simulations revealed a noticeable reduction in the aerofoil lift coefficient for the porous aerofoil. For unsteady solutions with a continuous or distinct series of vortices interacting with the aerofoil, the porosity showed insignificant damping of the lift coefficient amplitude. The porosity values investigated in the current exercise had indiscernible effect upon the unsteady lift-load alleviations caused by free stream disturbances.

On the basis of well-known literature, an analytical tool named LEAF (linear elastic analysis of fracture) was developed by the authors to predict the damage tolerance (DT) proprieties of aeronautical stiffened panels. The tool is based on the linear elastic fracture mechanics and the displacement compatibility method. By means of LEAF, an extensive parametric analysis of stiffened panels, representative of typical aeronautical constructions, was performed to provide meaningful design guidelines. The effects of riveted, integral and adhesively bonded stringers on the fatigue crack propagation performances of stiffened panels were investigated, as well as the crack retarder contribution using metallic straps (named doublers) bonded in the middle of the stringers bays. The effect of both perfectly bonded and partially debonded doublers was investigated as well. Adhesively bonded stiffeners showed the best DT properties in comparison with riveted and integral ones. A great reduction of the skin crack growth propagation rate can be achieved with the adoption of additional doublers bonded between the stringers.

Flying a light aeroplane involves a combination of pilot and aeroplane performing a set task, within a specific environment. The pilot is continuously sampling and selecting available sensory cues, interpreting those cues, making decisions and manipulating the primary controls (stick and rudder) to safely achieve flying objectives. The ‘feel’ of an aeroplane (a flying quality) is directly associated with the stick and rudder forces and how the aeroplane responds to control inputs. Classical theory has been applied to estimate the apparent (as felt by the pilot) longitudinal stick-free static stability (change of stick force with airspeed) of a typical, two-seat, high-wing light aeroplane. The theory has been extended to consider the effects of tail downwash and flap deflection. The results are compared with actual flight tests and show that the method may be used for the initial assessment of longitudinal stick-free static stability and more importantly, tendencies towards neutral or negative stability affecting flight safety.

This paper presents highlights from research conducted at the University of Liverpool to determine suitable fidelity criteria and guidelines for the modelling and simulation of the helicopter-ship dynamic interface environment. The paper begins by describing the characteristics of the helicopter-ship dynamic interface, explaining the motivation behind the research and reviewing the state-of-the-art in dynamic interface simulation. The development of a dynamic interface research environment based on an existing research simulator operated by the University of Liverpool is then described, before key results from a number of piloted simulation experiments are presented. These experiments were specifically designed to address fidelity sensitivity issues, such as, are unsteady airwake models necessary, or can a steady airwake model induce appropriate levels of pilot workload? What influence does the modelled ship geometry, or choice of atmospheric wind conditions have on the airwake model and on pilot workload? Finally, the paper concludes by briefly describing the relevance of these research findings to current and future industry programmes.

Current state-of-the-art trajectory prediction tools typically model aircraft as three-dimensional point-masses, and make a number of simplifying assumptions about the actual and anticipated dynamics states of the aircraft. They are typically based on predefined settings obtained from existing databases such as Eurocontrol’s Bada rather than real-time information, including on the environment, available onboard the aircraft. This significantly limits trajectory prediction performance. This paper proposes a high-accuracy four-dimensional trajectory prediction model for use onboard civil aircraft, as well as by ground-based systems, which addresses these limitations. It is designed for strategic traffic capacity optimisation and conflict-detection and resolution over time-horizons covering the entire duration of a flight. The model incorporates a number of features including a novel flight-control-system and an enhanced flight-script that incorporates new taxonomy and content thereby enabling better definition of aircraft intent. The accuracy of the model is characterised using operational data acquired during a real flight trial. Results show that the performance of the proposed model is significantly better than the current models. Its accuracy is better than the required navigation performance for departure, en route and Non-Precision-Approach phases of flight.

In this paper an associative-parametric approach is proposed in order to model the mesh of an aeronautical concept starting from a set of high-level structural primitives. This approach allows the designer to carry out the geometric modelling and the automatic mesh generation within one software environment in a fast and interactive way. The structural optimisation process is then simplified, with a relevant man-hours saving. A lower number of data transfers between different software is, moreover, involved with less problems related to the data corruption. To assure orders of continuity higher than C0 between adjacent instances, a suitable mathematical description of the structural primitives has been proposed. This description assures the maintenance of the required continuity constraints when the mesh is modified. Appropriate schemes of dependences are identified to guarantee the automatic propagation of the modifications complying with the continuity constraints.

In this paper a geometric sizing method for a small electric powered flying wing is proposed. The geometric sizing method aims to reduce the effects of variations in the power plant characteristics on endurance. This results in a single-objective design optimisation problem where the sensitivity to power plant characteristics of the endurance equation is minimised, constrained to Reynolds number, wing load, wing taper ratio, aircraft size and wing sweep angle. As a result, geometric characteristics of the flying wing such as span, tip chord and root chord are obtained. Flying wing aerodynamic characteristics are obtained by means of an inviscid fluid flow analysis program of the type low-order panel methods, known as CMARC. The optimisation problem involves a non convex function so that it is necessary to rely on heuristic programming methods. In particular an Evolutionary Algorithm based on differential evolution is considered.

Free-form deformation (FFD) is a method first introduced within the graphics industry to enable flexible deformation of geometric models. FFD uses an R3 to R3 mapping of a deformable space to the global Cartesian space to produce the geometry deformation. This method has been extensively used within the design optimisation field as a shape parameterisation technique. Typically it has been used to parameterise analysis meshes, where new design geometries are produced by deforming the original mesh. This method allows a concise set of design variables to be used while maintaining a flexible shape representation. However, if a computer aided design (CAD) model of the resulting geometry is required, reverse engineering techniques would need to be utilised to recreate the model from the deformed mesh. This paper extends the use of FFD within an optimisation routine by using FFD to directly parameterise a CAD geometry. Two methods of linking the FFD methods with the CATIA V5 CAD package are presented. Each CAD integration technique is then critiqued with respect to shape optimisation. Finally the set-up and initialisation of a case study is illustrated. The case study chosen is the aerodynamic optimisation of the wing-fuselage junction of a typical passenger aircraft.

Results from numerical simulations of the three dimensional flow in the intake-isolator of a dual mode scramjet are presented. The FANS calculations have utilised the SST k -ω turbulence model. The effect of cowl length and cowl convergence angle on the inlet mass capture ratio, flow distortion, shock strength and pressure rise are studied in detail. Three cowl lengths and four or five cowl convergence angles for each cowl length are considered. The predicted values of the dimensionless wall static pressure and inlet mass capture ratio are compared with experimental data reported in the literature. The numerical predictions are shown to agree well with the experimental data. In addition, details of the flow field such as shocks, expansion fans and shock boundary interaction are also captured accurately. Inlet unstart is also demonstrated for one case.

In various studies involving subsonic transition cavity flows, terms like ‘transitional-open’ and ‘transitional-closed’ cavity flow regimes are used in the literature. However, the exact fluid dynamic criteria to distinguish them have not been firmly established. The present work attempts to clarify this. Experiments to measure static pressure and fluctuating pressure distribution in a rectangular cavity with length-to-depth (L/D) ratio of 3 to 20 were performed. Based on pressure measurements on the front, bottom and rear faces of the cavity, additional criteria are established to clearly distinguish the range of critical L/D (or (L/D)cri) where these different transitional cavity flow regimes occur. The present work ascertains that the (L/D)cri for open and transitional-open cavity flow is 6·0-6·5, for transitional-open and transitional-closed cavity flow is 9-10 and for transitional-closed and closed cavity flow is 14-15. Flow visualisation qualitatively supports the flow pattern in the different cavity flow regimes.

The alleviation of gusts effects on a tiltrotor in aeroplane and helicopter operation modes obtained by an optimal control methodology based on the actuation of elevators, wing flaperons and swashplate is examined. An optimal observer for state estimate is included in the compensator synthesis, with the Kalman-Bucy filter applied in the presence of stochastic noise. Tiltrotor dynamics is simulated through an aeroelastic model that couples rigid-body motion with wing and proprotor structural dynamics. An extensive numerical investigation examines effectiveness and robustness of the applied control procedure, taking into account the action of both deterministic and stochastic vertical gusts. In addition, a passive pilot model is included in the aeroelastic loop and the corresponding effects on uncontrolled and controlled gust response are analysed.

This paper describes the computation of stress intensity factors (SIFs) for cracks in functionally graded materials (FGMs) using an extended element-free Galerkin (XEFG) method. The SIFs are extracted through the crack closure integral (CCI) with a local smoothing technique, non-equilibrium and incompatibility formulations of the interaction integral and the displacement method. The results for mode I and mixed mode case studies are presented and compared with those available in the literature. They are found to be in good agreement where the average absolute error for the CCI with local smoothing, despite its simplicity, yielded a high level of accuracy.