This paper describes a multidisciplinary robust optimisation framework for UAV conceptual design. An in-house configuration designer system is implemented to generate the full sets of configuration data for a well-developed advanced UAV analysis tool. A fully integrated configuration designer along with the UAV analysis tool ensures that full sets of configuration data are provided simultaneously while the UAV configuration changes during optimisation. The computational strategy for probabilistic analysis is proposed by implementing a central difference method and fitting distribution for a reduced number of Monte Carlo Simulation sampling points. The minimisation of a new robust design objective function helps to enhance the reliability while other UAV performance criteria are satisfied. In addition, the fully integrated process and a probabilistic analysis strategy method demonstrate a reduction in the probability of failure under noise factors without any noticeable increase in design turnaround time. The proposed robust optimisation framework for UAV conceptual design case study yields a more trustworthy prediction of the optimal configuration and is preferable to the traditional deterministic design approach. The high fidelity analysis ANSYS Fluent 13 is performed to demonstrate the accuracy of proposed framework on baseline, deterministic and RDO configuration.

Major results and findings of the numerical work package of the European high lift programme EUROLIFT are outlined. The main objective of these studies is to validate and test numerical methods for the prediction of high lift flows for transport aircraft configurations. The activities comprise the assessment of current CFD methods for 3D flows, evaluation of means for code improvement, and transition prediction. All aspects are especially devoted to high lift flow problems. A general capability to predict maximum lift on a simplified wing/fuselage high lift configuration is demonstrated by a variety of different numerical approaches. In general, major shortcomings are the reliability and the accurate simulation of large separation areas and the turn-around time to compute 3D lift polars. Advanced turbulence modelling and numerical solver features, such as the preconditioning technique, show a potential to overcome these deficiencies. Promising results with respect to transition prediction were obtained on a swept high lift wing using a database method. The results obtained in the numerical activities represent major ingredients on the way to a consistent numerical approach for the simulation of transport aircraft high lift configurations including all maximum lift determining effects.

An analysis of an extensive series of unsteady and steady pressure measurements at transonic speeds on a rigid 12% thick NACA 16 series aerofoil reveals a useful classification of the type of excitation characteristics of particular flow regions.

The pressure fluctuation measurements suggest that when the shock from the aerofoil reaches the tunnel roof, serious interference (both dynamic and static) occurs.

The RAE 2 ft × 1½ ft Transonic Tunnel used for the tests has low levels of flow unsteadiness at low frequencies, making it a useful facility for basic research in time-dependent aerodynamics.

An approximate theory is described which takes into account the influence of the wake on the aerodynamic derivatives of an oscillating rotor blade. It is shown that the most important parameter is the number of cycles of the oscillation which occur during one revolution of the rotor. For integral values of this parameter the flexural damping is shown to become very small at small angles of incidence. The flutter characteristics of a simple rotor blade are calculated on the basis of this theory, and it is found that the wake causes radical changes in the critical speeds. The theory is in good qualitative agreement with experiment.

A procedure for developing efficient aeroelastic reduced order models (ROMs) for aerospace structures containing geometric nonlinearities is described. The structural modelling is based upon a combined modal/FE approach that describes the non-linear stiffening effects from results of non-linear static analyses for a range of prescribed inputs. Once the structural ROM has been defined, it is coupled to the rational fraction approximation of the doublet lattice aerodynamic model corresponding to the wing planform. The aeroelastic model can then be used to predict the dynamic aeroelastic behaviour of the defined structure. The methodology is demonstrated on the aeroelastic model of a flexible high aspect ratio wing with the static deflections, LCO behaviour and gust response being predicted.

This paper presents a systematic formulation for the simulation of rigid body dynamics, including the short period dynamics, inherent to stage separation and jettisoning parts of a satellite launcher. This also gives a review of various types of separations involved in a launch vehicle. The problem is sufficiently large and complex; the methodology involves iterations at successively lower levels of abstraction. The best choice to tackle such problems is to use state-of-the-art programming technique known as object oriented programming. The necessary classes have been identified to represent various entities in the launch vehicle separation process (e.g., gravity, aerodynamics, propulsion and separation mechanisms etc.). Simple linkages are modelled with suitable objects. This approach helps the designer to simulate a launch vehicle separation dynamics and also to analyse separation system performance. To examine the influence of the design variables on the separating bodies, statistical analyses have been performed on the upper stage separation process and pull out of ongoing stage nozzle from the spent stage of a multistage rocket carrier using retro rockets.

The matrix methods of structural analysis developed specifically for use on modern digital computers have now become universally accepted in structural design. These methods provide a means for rapid and accurate stress and deflection analysis of complex structures under static and dynamic loading conditions and they can also be used very effectively for the stability analysis. In the conventional stability analysis two possible approaches are normally used; either the differential equations describing the structural deflections are formulated and the lowest eigenvalue representing the buckling load condition is found for a given set of boundary conditions, or alternatively, if the differential equations are too difficult to prescribe, approximate deflection shapes are used in the strain energy expression for large deflections which is subsequently minimised, leading to the stability determinant whose lowest root represents the instability condition. When designing complex structures the conventional methods of finding buckling load conditions are extremely difficult to apply, and therefore in such cases we have to rely on the matrix methods of stability analysis.

This paper describes the development and validation of a high fidelity simulation model of the Bell 412 helicopter for handling qualities and flight control investigations. The base-line model features a rigid, articulated blade-element formulation of the main rotor, with flap and lag degrees of freedom. The Bell 412 HP engine/governor dynamics are represented by a second-order system. Other key features of the base-line model include a finite-state dynamic inflow model and lag damper dynamics. The base-line model gives excellent agreement with flight-test data over the speed range 15-120kt for on-axis responses. Prediction of off-axis responses is less accurate. Several model enhancement options were introduced to obtain an improved off-axis response. It is shown that the pitch/roll off-axis responses in transient manoeuvres can be improved significantly by including wake geometry distortion effects in the Peters-He finite-state dynamic inflow model.

The aerodynamic interference between typical tanker and large receiver aircraft during air-to-air refuelling was modelled using a simple horseshoe vortex representation of the tanker wake whilst the aerodynamic forces and moments acting on the receiver were calculated using the vortex lattice method. For each combination of tanker and receiver, the aerodynamic loads depend mainly on the vertical separation, whilst the receiver tailplane position on the fin strongly effects the pitching moment acting on the receiver due to the tanker downwash. Particular attention was given to a high T-tail receiver aircraft. The predicted increase in the receiver drag compares favourably with available flight test data. Changes in trim in pitch of the receiver were analysed and related to the static stability of the aircraft. Aerodynamic derivatives due to the position and attitude of the receiver aircraft within the tanker downwash field were calculated and solutions of the linearised equations of motion obtained for the high T-tail receiver refuelling from different tanker aircraft at different flight conditions. Typically the receiver exhibited a divergent oscillation at the normal refuelling position. The dynamic stability characteristics are consistent with the trends observed in flight tests on the tendency of the receiver to oscillate in pitch.

With the return of British civil aviation to normal operating conditions after the war it was found that several problems which had previously given little cause for concern had become acute. More especially, the increase in the volume of traffic, particularly under instrument flying conditions, together with certain indirect consequences of the greater sizes and weights of aircraft, had brought about greatly increased congestion around major airports, and there was, therefore, an urgent requirement for an effective system of close air traffic control to ensure the safety and rapid movement of the traffic.

The Ministry of Civil Aviation immediately gave much attention to this problem of air traffic control, but at once found itself in serious difficulties on account of lack of sufficiently precise knowledge of the traffic and the traffic pattern. The then Controller of Technical and. Operational Services, Sir Conrad Collier, realised that the obtaining of the necessary data concerning air traffic control and other matters of technical interest was essentially a matter of operational research, and in 1947 a small operational research section was established which began work in the autumn of that year.

Under zero-gravity conditions, the liquid surface in a container has a displaced equilibrium position due to surface tension effects; this depends upon the magnitude of the contact angle to the wall. Free and forced sloshing have been investigated for these conditions. It was found that the natural frequencies of the liquid increases with the contact angle to a maximum value at a contact angle slightly above π/2 and then decreases. This behaviour is not symmetrical within the wetting and non-wetting range. Thus, a completely non-wetting liquid has larger natural frequencies than a completely wetting liquid. It was also found that the sloshing mass of a wetting liquid is considerably larger under zero-gravity than under normal gravity conditions and that a non-wetting liquid exhibits substantially reduced sloshing mass, indicating that the influence of liquids with a small contact angle on the dynamics of a vehicle with liquid containers (with a free liquid surface) may be quite pronounced.

In this paper an approach to fault-tolerant flight control system design based on the integration of sensor/actuator fault detection and isolation (FDI) and reconfigurable control is presented. The effects of the sensor and actuator faults in the innovation process of the Kalman filter are investigated, and a decision approach to isolate the sensor and actuator faults is proposed.

The reconfigurable control algorithm based on the Extended Kalman Filter (EKF) is presented. The control reconfiguration procedure is executed by considering the identified control distribution matrix. In the simulations, the longitudinal dynamics of an aircraft control system are considered, and control reconfiguration is examined. In the paper, the principal block diagram of a fault tolerant aircraft control system is offered.

The ability to track time varying frequency and damping parameters using on-line versions of seven time domain system identification algorithms; Least Squares, Double Least Squares, Correlation Fit, Instrumental Variables, Instrumental Matrix with Delayed Observations, Extended Least Squares and Maximum Likelihood, is examined. Comparisons are made on results obtained from data generated from various simulated time varying systems, corrupted with both input and measurement noise, as a preliminary step before advancing onto real data. Only the Maximum Likelihood and Double Least Squares methods gave acceptable results on the most difficult data sets that were considered. The choice of an exponential weighting factor is crucial to the performance of the on-line methods when applied to data from time varying systems. The effect of assigning various values of the weighting factor is demonstrated. Further work is required to determine the optimum schemes for defining this parameter.

Nowhere in the literature of helicopter noise can one find a general understanding; only isolated, inconsistent, and often inadequately documented facts. These could, of course, be reviewed but hardly with profit; it is easier to ask critical questions of nature than of previous investigators.

For fundamental reasons the helicopter in its present state of development is not in fact matched to the bulk, low-cost transport of passengers. But for equally good reasons the phenomena of aerodynamically generated noise now bid fair to banish from inhabited localities all other forms of vertical take-off and landing aircraft so that, if a city-centre to city-centre aircraft transport system is ever to be practicable, it will most probably be based on some metamorphosis of the helicopter.

Two numerical methods are presented for the computation of steady and unsteady Euler flows. These are applied to steady and unsteady flows about the NACA 0012 aerofoil, using structured grids generated by the transfinite interpolation technique. An explicit central-difference scheme is produced based on the cell-vertex method of Ni modified by Hall. The method is second-order accurate in time and space, and with flow quantities stored at boundaries the boundary conditions are simple to apply. This is a definite advantage over the cell-centred approach of Jameson, where extrapolation of the flow quantities is required at the boundaries, making unsteady boundary conditions difficult to apply. An explicit upwind-biased scheme is also produced, based on the flux-vector splitting of van Leer. The method adopts a three stage Runge-Kutta time-stepping scheme and a high-order spatial discretisation which is formally third-order accurate for one-dimensional calculations. The upwind scheme is shown to be slightly more accurate than the central-difference scheme for steady aerofoil flows, but it is not clear which is the more accurate for unsteady aerofoil flows. However, the central-difference scheme requires less than half the CPU time of the upwind-difference scheme, and hence is attractive, especially when considering three-dimensional flows. The transfinite interpolation technique is ideal for generating moving structured grids due to its simplicity, and grid speeds are available algebraically by the same interpolation as grid points. The method is also ideal for use in a multi-block approach.

In seeking to analyse the stress distribution in a complex structure, our object is really twofold. In the first place, if the stress in each part is accurately known, the strength of each part separately can be assessed on the basis of tests on simple units and the necessity for testing the complete structure may be avoided. In the second place, if the stress distribution can be expressed as a function of the geometry of the structure, the effect of varying that geometry, particularly with the object of realising Holmes’ one-hoss-shay, can be examined theoretically and arbitrary-features of the design can be eliminated. If our knowledge were complete, any structure designed to a maximal or minimal condition would not permit any variation from one ideal form, the design would be uniquely determined by the stated requirements.

In this article a magnetic suspension and balance system suitable for wind tunnel application is discussed. General considerations are presented that illustrate the nature of the problems to be solved as well as one solution of these problems.Finally some initial wind tunnel data are presented.

The paper presents the results of the aerodynamic optimisation of an Unmanned Aerial System's wing using a morphing approach. The shape deformation of the wing is achieved by placing actuator lines at several positions along its span. For each flight condition, the optimal displacements are found by using a combination of the new Artificial Bee Colony algorithm and a classical gradient-based search routine. The wing aerodynamic characteristics are calculated with an efficient nonlinear lifting line method coupled with a two-dimensional viscous flow solver. The optimisations are performed at angles of attack below the maximum lift angle, with the aim of improving the Hydra Technologies UAS-S4 wing lift-to-drag ratio. Several configurations of the morphing wing are proposed, each with a different number of actuation lines, and the improvements obtained by these configurations are analysed and compared.

A brief narration on significant accomplishments in computational fluid dynamics (CFD) for basic research and aerospace application is attempted to highlight the outstanding achievements by scientists and engineers of this discipline. To traverse such a vast domain, numerous and excellent contributions to CFD will be unintentionally overlooked by the author’s limited exposure. Nevertheless it is an ardent hope that the present abridged literature review will aid to reaffirm excellence in research and to identify knowledge shortfalls both in aerodynamics and its modeling and simulation capability. The future modeling and simulation technology needs, as well as potential and fertile research areas, are humbly put forth for consideration.

Non-linear behaviour is usually undesirable for the operation of any system and it needs to be minimised in flight control systems in order to provide the pilot with a predictable and well-behaved aircraft. The non-linear response characteristics exhibited by all aircraft and their flight control systems can be characterised by using nonlinear functions to simulate and analyse their behaviour. This paper explores the types of physical non-linearities that exist in aircraft flight control systems and provides some examples. It describes how the non-linear stability might be analysed and addresses the subject of actuation systems modelling. The final section provides recommendations for minimising non-linear behaviour and gives some general guidelines on how to deal with non-linear characteristics.