Many situations arise in which aerospace structures are subjected to high frequency excitation, in the sense that the wavelength of the induced dynamic response is much shorter than the overall dimensions of the structure. The application of the conventional finite element method to this type of problem faces two difficulties: (i) the short wavelength of the structural deformation requires the use of many elements, which renders the method computationally expensive or even impracticable, and (ii) the response of the structure at high frequencies can be very sensitive to structural detail, and thus response predictions for an ‘ideal’ structure may differ significantly from the performance of the actual system. For a number of years research effort has been directed towards the development of alternative analysis methods for high frequency vibrations, and recent developments in this area are reviewed in the present paper. The methods considered are: (i) hierarchical versions of the finite element method, (ii) the dynamic stiffness method, (iii)periodic structure theory, (iv)statistical energy analysis and (v) wave intensity analysis.

An experimental investigation to determine the effectiveness of a wing-tip device designed to attenuate the trailing vortex of a lifting, flat-plate wing is detailed. The induced-breakdown vortex attenuation device (IBVAD) is essentially a toed-in, vertically-oriented sheared tip. Results are presented comprising force-balance, water tunnel flow visualisation and wake studies using a seven hole probe. The data shows that compared with an equal span rectangular wing, variations of the IBVAD device can reduce drag at moderate-to-high lift coefficients, as well as reduce die maximum rotary velocity of the trailing vortex by 28%. Axial vorticity in the wake vortex was significantly attenuated compared with the rectangular wing, with a 52·6% reduction in the magnitude of die maximum vorticity. The vortex core radius was also seen to be larger than die rectangular wing. However, even substantial attenuation and redistribution of trailing vorticity may be insufficient to reduce the wake hazard.

Inviscid and viscous computational models have been used to predict two-dimensional steady separated flow around an aerofoil with a spoiler. First, an attempt is made to construct an inviscid panel method in which piece-wise linear vortex panels are placed on the aerofoil and spoiler. The separated region is modelled with potential flow analysis by using free vortex sheets placed on the separation streamline from the spoiler tip and aerofoil trailing-edge. The shapes of the vortex sheets require an iteration to be established. The calculation is based on an assumption of the length of the free vortex sheets. Second, a viscous procedure based on a finite control volume scheme is used to solve the Navier-Stokes equations for turbulent flow by making used of the k-ε model. An algebraic pressure correction is incorporated in the calculation. The predictions from the two models are in reasonable agreement with experimental measurements.

The vorticity production by a single vortex generator jet in a turbulent boundary layer at high speed and zero pressure gradient was studied by analysing surface-flow patterns. The investigation involved a range of jet mass flow rates and skew angles while the freestream Mach number and jet pitch angle were held constant. A qualitative analysis of the skin friction patterns demonstrated the effective role of the skew angle as a bifurcation parameter. The vortex production is classified into three main regimes as the skew angle increases from zero (fully downstream) to 180° (fully upstream). Better understanding of the physical mechanism behind the transition from the first regime to the second is crucial in the design of the airjet vortex generator.

A moving mesh method for the computation of compressible viscous flow past deforming and moving aerofoils is described. It is based on a well established time-marching finite-volume scheme, which has been widely used for steady compressible flows. An implicit real-time discretisation is used and the equations are integrated via a dual-time scheme. This involves the introduction of derivatives of a fictitious pseudo time. The solution at each real-time step involves seeking solutions which are steady-state solutions in pseudo time. This approach decouples the stability and accuracy limitations of the scheme and permits large real-time steps to be chosen. Also well-proven convergence acceleration techniques developed for steady flows such as local-time stepping, residual averaging and multigrid may be used in the pseudo-time stepping scheme without compromising real-time accuracy. A sequence of body-conforming grids and corresponding grid speeds is required, where the inner and outer boundaries of the grid move independently. The required grids and speeds are found using a transfinite interpolation technique. Applications of the method to external compressible flows are shown, including results from a parallel version of the computer program.

Results obtained from measurements of the dynamic induced velocity field beneath a model rotor are presented. The collective and cyclic pitch of a four bladed rotor of 1·54 m diameter were excited at frequencies up to 1·5 times rotor shaft speed. Flow measurements were taken using hot wire anemometry probes and a laser doppler anemometer. A range of radial and vertical positions near the rotor disc were investigated.

Analysis of the dynamic induced flow response is conducted in the frequency domain and it is found that there are significant radial and azimuthal variations which depend on the frequency of excitation. It is also observed that a change in the character of the inflow response occurs near and above the shaft rotational frequency and that vertical measuring distance from the rotor significantly affects the measured responses.

Some results for the case of cyclic excitation are given. These show, that contrary to momentum theory predictions, the highest induced velocities in the dynamic case do not occur over the area of the disc where the blade pitch is at its maximum.

Overall, the results show that the dynamic induced velocity field is highly complex and heavily influenced by the distribution of time varying shed vorticity within the wake.

Simulated hailstones were made to impact on the rotating spinner and fan assembly of a Williams FJ44 engine. The mass distribution of ice behind the fan was determined by use of a suction tube technique. Suction was added to ensure that the tube did not affect the flow through the fan assembly. The strong air flow behind the fan meant that the ice caught by the tube melted and evaporated. This made it difficult to accurately determine the mass distribution of ice. As a result, the simulated hailstones were made from a water-salt solution so that the weight of salt residue could be measured after the water had evaporated, and hence the amount of ice caught was determined. A parametric study into the hail ingestion characteristics of the fan assembly was carried out. The parameters studied included the radial position of the impact point, the rotational speed of the fan and the position of the splitter between the core engine and bypass duct. The results showed that the impact position had a major effect on the overall ice distribution, which was determined by the combination of blade geometry at the impact point and the rotational speed. The splitter position was shown to have a significant effect on the amount of ice passing into the bypass duct.

The purpose of a joint is to transfer load between the two parts being joined. As a result of this load transfer there will be a stress variation in the components in the joint region, as well as stresses in the joining medium (fasteners or adhesive). In fhjs work, the stress results of a two-dimensional finite element (FE) analysis are used to understand failure modes of a bolted joint in low-temperature cure carbon fibre reinforced plastic (CFRP) woven laminates loaded in tension, and to predict the bearing strength. Good agreement with experimentally observed damage modes and strengths is achieved with, in some cases, a difference of less than 10%.

The methods of non-linear dynamics are applied to the longitudinal motion of a vectored thrust aircraft, in particular the behaviour at high angles-of-attack. The model contains analytic non-linear aerodynamic coefficients based on Nasa windtunnel tests on the F-18 high alpha research vehicle (Harv). The equilibrium surfaces are plotted against thrust magnitude and thrust deflection and are used to explain the behaviour. When the aircraft is forced with small thrust deflections whilst in poststall equilibrium, chaotic motion is observed at certain frequencies. At other frequencies, several limiting states coexist, e.g. a chaotic attractor and a limit cycle, or two limit cycles. The steady state behaviour then depends on the initial conditions. The non-linear pitching moments are shown to have a significant effect upon the model. Omitting these terms, the region of chaos shrinks, this is not observed for lift.

A low Reynolds number turbulent boundary layer is subjected to a relatively strong rate of suction, which is applied through a short porous strip. Downstream of the strip, the effectiveness of a riblet surface is assessed by comparison,to that of a smooth wall. Measured mean velocity and longitudinal turbulence intensity profiles indicate that the riblets do not hinder the action the suction can have on the flow.

Presented in this paper are the preliminary results from a blade-vortex interaction (BVI) test series conducted in the University of Glasgow's 1·61m x 2·13m closed-return, low-speed windtunnel, in which a single vortex interaction with a rigid, loaded rotor blade was studied. This work is an extension of that by Horner and Galbraith, which describes the flow phenomena present during single vortex interaction with an unloaded blade. The paper presents blade surface pressure data recorded by 72 pressure transducers located in the outer regions of the blade during these interactions. Integrated Cn and Cm1/4 data are also presented to provide a detailed history of a BVI event. In addition, a method of isolating the bound circulation and vortex induced effects is presented, from which it is concluded that for a BVI event the vortex-induced effects are (to a first approximation) independent of rotor pitch setting.

The integration inverse method has been extensively investigated in the past decade; however, none of the researches fully addresses the stability analysis of the method that is crucial to successful implementation. This paper presents a stability test to analyse the global stability of the integration inverse method for linear time-invariant systems. A stable solution may be obtained from careful selection of the discretisation interval using the proposed stability test. A discrete model is derived to approximate the Newton's scheme in the inverse method. With this approximate model, the stability of the inverse method can be examined. The stability test is exact for linear systems and can be extended to the inverse method for non-linear aircraft simulations by considering an equivalent linear model for each point of the flight envelopes. Guidelines for selection of appropriate reference points in the inverse simulation are given.

A high order boundary element formulation is presented and applied to the solution of potential, incompressible flows around non-lifting and lifting configurations. The high order numerical algorithm is based on a bicubic interpolation of both geometry and quantities over each element of discretisation of the boundary. Numerical validation of the formulation is performed by studying the aerodynamic solution around fuselages and wings, and making comparisons with existing numerical results.