A method is described for obtaining the displacements (and hence the stresses) of a linear elastic system under a given transient load in a more accurate and more expeditious manner than is possible by the standard method. For simplicity, ho account is taken of damping forces in describing the method.

The method is of interest to engineers and others who require to obtain numerical answers to problems of the type described above. It depends upon the fact that in many practical cases of transient loads the inertia forces in the series of normal modes of an elastic system converge rapidly, compared with the series of “normal” loadings in terms of which the spatial distribution of the applied load can be expressed.

This note attempts to clarify some basic ideas, and correct some common misunderstandings, on wing flexure-torsion flutter. Topics include:

• (i) an understanding of the physics of wing flexure-torsion flutter through a number of specific examples,

• (ii) an examination of the energy interpretation of the physics of flutter, indicating its limited usefulness,

• (iii) a discussion of the significant differences between wing flexure-torsion flutter and the Duncan flutter engine,

• (iv) a graphical representation to assess the contribution of various parameter to flutter onset,

• (v) subcritical response below the critical flutter speed,

• (vi) some practical wing flutter considerations.

Static pressures have been measured in the separated flow region behind backward-facing steps in a wind tunnel, suitable corrections being applied for the model blockage. The results indicate a reasonably well defined similarity in the pressure distribution.

This paper presents a general procedure for calculating the natural frequencies of rectangular plates continuous over identical and equally spaced elastic beams which are simply-supported at their ends. Arbitrary boundary conditions are permissible on the other two edges of the plate. The results are compared with those obtained by using the orthotropic plate approximation for the system

The unsteady air flow over and around the helicopter landing deck of a naval vessel is known to cause high pilot workload and to limit the helicopter's operational envelope for launch and recovery. Previous research has suggested that modifications to the ship's hangar edges can beneficially modify the flow over the deck. This paper examines the effectiveness of five hangar-edge modifications using computational fluid dynamics–generated airwakes and flight mechanics modelling, as well as piloted flight trials in a motion-base simulator. Results are presented, in terms of unsteady helicopter loads and pilot workload ratings, for modifications to the windward vertical rear edge of the hangar and with an oblique wind. The results demonstrate that while the airwake can be altered by superstructure modifications, the integrated effect of the altered airwake on the entire helicopter does not necessarily give the desired result; indeed of the five modifications tested, two were seen to be beneficial while three caused an increase in pilot workload compared with the unmodified hangar. Overall, the paper shows that the airwake can be modified by superstructure design changes, and that the effect on the helicopter can be measured through modelling and simulation. It is also demonstrated that making judgements on the severity of the airwake based on the aerodynamic flow field alone can be misleading. The benefit of these simulation tools is that they can be used during the ship design process to evaluate the effect of the superstructure aerodynamics, rather than wait until after the ship is built.

The challenges of helicopter simulation are being tackled across a broad front as technology is developed to meet the needs of Industry. Traditionally, the strongest impetus has come from the training community and this is likely to continue for some time as simulation technology advances at increasing pace, raising fidelity standards. The development of PC-based simulation technologies is providing a significant spur in this development and lowering the cost, making complete simulation systems of reasonably high fidelity available to smaller organisations. This paper describes the first year of operation with such a system at the University of Liverpool – HELIFLIGHT. With its full motion, wide field-of-view visuals, programmable force-feel system and the comprehensive FLIGHTLAB modelling environment, we describe the HELIFLIGHT system as high fidelity and the first year of utilisation saw extensive use in a variety of handling qualities and pilot-vehicle technology research and teaching. Some of these are described in the paper and from a year of many highlights, the EU-funded programme to develop handling qualities for a civil tilt rotor aircraft is selected to demonstrate the versatility of the HELIFLIGHT system.

An incompressible three-dimensional turbulent wall jet originating from a circular orifice located adjacent to a plane wall is studied both theoretically and experimentally. An approximate similarity analysis predicts that the two transverse length scales, l0 and L0, and the inverse of the mean velocity scale grow linearly with distance downstream x from the orifice. Experimental measurements of mean velocity and longitudinal turbulence intensity profiles were made both in air and water with hot-wire and hot-film anemometers respectively. The behaviour predicted by the similarity analysis was verified. It was found that the rate of growth of the length scale normal to the plane wall, dl0/dx, was somewhat less than that found for a two-dimensional wall jet, whereas the rate of growth of the length scale in the lateral direction, dL0/dx, was about seven times greater than dl0/dx.

A series of mixed-compression hypersonic inlets at different bleeding rates were simulated at different freestream conditions in this paper. The unstart/restart characteristics of hypersonic inlets were analysed and the reasons why the unstart/restart phenomenon is in existence is presented. The unstart/restart characteristics of hypersonic inlets at different bleeding rates were given. The effects of boundary-layer bleeding on the performance parameter (mass-captured coefficient, total-pressure recovery coefficient), starting and restarting Mach number of hypersonic inlets were discussed. In conclusion, boundary-layer bleeding can improve the performance parameter of hypersonic inlets, and can reduce the starting and restarting Mach number, and can broad the operation range of the hypersonic inlet.

An experimental investigation of the viscous decay of a steady, three-dimensional, helical vortex is presented. The vortex was generated by a rectangular, symmetrical half aerofoil (N.A.C.A. 0009) of 10·5 in. span, 3 in. chord, cantilevered from the wall of the circular test section of a low speed wind tunnel. Only the local total head and flow directions were measured with a five-hole pressure sensitive probe at one to eight chord lengths behind the trailing edge of the aerofoil, and the radial, tangential and axial velocity distributions in the trailing vortex were derived. The static pressure variation in the vortex flow and the difficulties in its accurate measurement are discussed. All tests were performed at a Reynolds number based on the free-stream velocity and chord length of the aerofoil of 10,000.

The various experimental velocity distributions in the vortex were in good agreement with a linearised, axi-symmetric, laminar, incompressible, viscous vortex due to Newman. The decay of velocity and geometry parameters, however, were found to be approximately eight to ten times faster than predicted by the analysis based on a laminar kinematic viscosity. It was also observed that the amount of circulation which actually rolled up in the trailing vortex was only about 58 per cent of that expected around the (N.A.C.A. 0009) aerofoil used for this investigation.

Wire gauzes are used in air duct systems and wind tunnels for many purposes. Methods of estimating the pressure loss through a gauze currently available are often found in practice to be unsatisfactory. In this paper data obtained by several experimenters are correlated. It is shown that the effect of Reynolds number on pressure drop is similar for a large range of gauze porosities. An empirical method of estimation of the pressure drop is given which can be used with confidence for gauzes of porosity (defined as the proportion of area not blocked by the wires) from 0·27 to 1·00, and over a range of Reynolds number, based on approach velocity and wire diameter, from 20 to 600.

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.

Inverse simulation is becoming a more widely used technique in flight mechanics studies. The ability to predict control and response time histories for a predefined manoeuvre or task has obvious benefits in certain applications (handling qualities studies related to the ADS-33 requirements for helicopters, for example). The main criticisms of the technique have always been concerned with the numerical problems usually encountered when solving the equations of motion in the inverse manner. In this paper these problems are highlighted and a simple technique which overcomes them is presented. This technique provides a robust mechanism whereby reliable and consistent inverse simulations are possible.

This paper examines the dynamic stalling of a finite wing of aspect ratio 3·0 when subject to constant pitch motions up to and beyond stall. In particular, unsteady surface pressure data were obtained at 192 locations on the wing surface and these were then analysed to provide information on the nature and phasing of dynamic stall events in both the chordwise and spanwise directions. It was also possible to obtain sectional force and moment coefficients by integration of the pressures measured on specific chordal arrays. This provided valuable insight into the load distribution on the wing throughout the range of motion. On this basis, it was established that the wing loading distribution was consistent with conventional understanding of steady wing loading up to the incidence at which the dynamic stall vortex was initiated. Beyond this point, the formation and subsequent convection of the vortex structure was found to be strongly three-dimensional but, nevertheless, exhibited many of the features of two-dimensional dynamic stall.

An Euler optimisation for a BWB configuration with winglets incorporating an array of three-dimensional shock control bumps is carried out by employing an efficient adjoint-based optimisation methodology. A high fidelity multi-block grid with over two million grid points is generated to resolve the shape of the 3D shock control bumps, the winglet as well as the overall BWB shape, which are parameterised by over 650 design variables. In order to perform such a large aerodynamic optimisation problem feasibly, the optimisation tools such as the flow solver and the adjoint solver have to be parallelised with a good parallel efficiency. This paper reports the parallel implementation efforts on the adjoint solver; especially on the calculation of the sensitivity derivatives, which has to be looped over the total number of design variables. Results from the optimisation of the wing master sections, winglet aerofoil sections and the three dimensional bumps indicate a significant improvement regarding the aerodynamic performance against the baseline geometry for the given planform layout of the aircraft.

The buffet characteristics of a supercritical aerofoil were investigated in the High Reynolds Number Two-Dimensional Test Facility of the National Aeronautical Establishment (NAE). The test was performed quite deep into the buffet régime. The buffet onset boundary was determined from the divergence of the unsteady balance normal force. Steady and unsteady pressure measurements were obtained for shock-induced separation with reattachment as well as fully separated flows. For flow conditions where discrete shock wave oscillations occurred, ensemble-averaging of the unsteady pressures determined the propagation velocity of the pressure wave induced by periodic shock motion. A model of the self-sustained oscillatory shock motion due to shock-boundary layer interaction was formulated. Broad-band cross-correlations of the pressure field determined the downstream convection velocities of the turbulent eddies of separated flows.

This paper addresses the development of multiple UAV deployment simulation models that include representative aerodynamic cross-coupling effects. Applications may include simulations of autonomous aerial refuelling and formation flying scenarios. A novel wake vortex model has been developed and successfully integrated within a Matlab/Simulink simulation environment. The wake vortex model is both sufficiently representative to support studies of aerodynamic interaction between multiple air vehicles, and straightforward enough to be used within real time or near real time air-to-air simulations. The model integration process is described, and computational results of a two-vehicle-formation flight are presented.

A systematic methodology for the thermodynamic optimisation of civil bypass engines (turbofan or advanced propulsors) is presented, which would be useful for designing air-breathing engines based on “clean-sheet analysis”. The process starts with establishing an optimum specific thrust for the engine based on an economic analysis (installation constraints, noise regulations etc. also need to be considered). The task of the optimisation process is then to find the combination of optimum values of fan pressure ratio, overall pressure ratio, bypass ratio and turbine entry temperature concurrently that maximises overall efficiency at the fixed specific thrust. This procedure is quite different from the usual single-variable parametric performance studies which do not give proper optimum values and may involve large excursion in the value of the specific thrust unacceptable for a particular mission. Additionally, several, simple and explicit, analytical relations are derived here from fundamental principles, which perform well against numerical optimisation performed by a specialist computer program employing iterative and advanced search techniques. The analytical relations accelerate the optimisation process and offer physical insight. Present numerical computations with real gas properties have established new concepts in turbofan optimisation (for example, the existence of an optimum bypass ratio and optimum turbine entry temperature). The question of optimum jet velocity has been addressed. An analytical expression for the optimum jet velocity at a given bypass ratio has been derived which performs well against numerical optimisation results.

A new wing-tip concept with morphing upper surface and interchangeable conventional and morphing ailerons was designed, manufactured, bench and wind-tunnel tested. The development of this wing-tip model was performed in the frame of an international CRIAQ project, and the purpose was to demonstrate the wing upper surface and aileron morphing capabilities in improving the wing-tip aerodynamic performances. During numerical optimisation with ‘in-house’ genetic algorithm software, and during wind-tunnel experimental tests, it was demonstrated that the air-flow laminarity over the wing skin was promoted, and the laminar flow was extended with up to 9% of the chord. Drag coefficient reduction of up to 9% was obtained when the morphing aileron was introduced.

The forces on an aerofoil have been measured at various angles of incidence for 0.5 < R < 2,000.

Below R = 2 the lift is found to be given approximately by :—

L, = (3/2) µlV sin 2x

and the drag at all angles of incidence tends to:

D = µlV.

An approximate theoretical estimate is also given of the effect of spacing, on the drag of a series of flat plates at R → o.

The aim of the study described herein was to develop and verify an efficient neural network based method for extracting aircraft stability and control derivatives from real flight data using feed-forward neural networks. The proposed method (Modified Delta method) draws its inspiration from feed forward neural network based the Delta method for estimating stability and control derivatives. The neural network is trained using differential variation of aircraft motion/control variables and coefficients as the network inputs and outputs respectively. For the purpose of parameter estimation, the trained neural network is presented with a suitably modified input file and the corresponding predicted output file of aerodynamic coefficients is obtained. An appropriate interpretation and manipulation of such input-output files yields the estimates of the parameter. The method is validated first on the simulated flight data using various combinations and types of real-flight control inputs and then on real flight data. A new technique is also proposed for validating the estimated parameters using feed-forward neural networks.