In the era of Unmanned Aerial Systems (UAS), an onboard autopilot occupies a prominent place and is inevitable for many of their modern applications. The efficacy of autopilot heavily relies upon the accuracy of the sensors employed and the capability of the onboard flight controller. In general, aerodynamic behaviour and flight dynamic capabilities of Unmanned Aerial Vehicles (UAVs) govern the selection and the design of flight controllers. Precise modeling of linear aerodynamic characteristics from flight data can be achieved using many of the existing classical parameter estimation techniques such as Output Error Method (OEM), Equation Error Method (EEM), and Filter Error Method (FEM). However, all the classical methods may not be readily applicable for aerodynamic modeling in nonlinear flight envelopes. The current manuscript is an attempt to exploit the capabilities of the Artificial Intelligence (AI) technique, named Particle Swarm Optimisation (PSO), in combination with Least Squares (LS) cost function to perform linear as well as nonlinear aerodynamic parameter estimation. The aforementioned task is accomplished by considering flight data from manoeuvers pertaining to linear angles of attack, moderate and near stall flight envelopes of two different UAVs with cropped delta planform geometry. Parameters estimated using the proposed LS-PSO method are consistent with minimum standard deviation and are on a par with OEM estimates. The proposed LS-PSO method enhances the capabilities of LS-based EEM while estimating stall characteristic parameters, which was not possible with LS alone. The longitudinal and lateral-directional static parameters estimated from the full-scale wind tunnel testing of the two UAVs were also used to corroborate the results obtained from the flight data using the LS-PSO method.

The wholly supersonic flow past a two-dimensional wedge is analysed on the assumption that release of chemical energy into the stream can be accomplished across a thin discontinuous plane flame front attached to the apex. Forces experienced by the wedge are calculated and representative flow patterns exhibited. Some typical interactions between the flame and shocks or centred simple waves are discussed, with emphasis on the use of pressure-flow-deflection diagrams to obtain results.

The inverse problem of determining the supersonic flow past a thin symmetrical wing at zero incidence with given surface pressure distribution is solved for wings of arbitrary plan form. Expressions are obtained for the vertical derivative of the potential on the wing surface from which, using the linearised boundary condition of zero normal velocity at the surface, the profile of the wing can be designed. The integral equation method adopted by J. C. Evvard and extended by G. N. Ward is used. The analysis cannot be applied to pointed wings, whose leading edges are subsonic. The results in Part I are completely general and are applied to specific problems in Part II.

The CHAIRMAN, in introducing the Author, said that Dr ROBERTS was very well known by everybody in the helicopter field m this country and abroad He had taken his B Sc Engineering Degree at Queen Mary College and his Ph D at Imperial College He had served with General Aircraft Ltd, and had also been a lecturer at Imperial College Subsequent to this appointment he had been with Saunders-Roe as Chief Project Engineer (Helicopters), and was now on the staff of Fairey Aviation where initially he had been Chief Aerodynamicist and was now Chief Project Engineer Dr Roberts was a Founder Member and a member of the Council of the Association

A general procedure for the analysis of thermal stresses in a polygonal elastic plate due to heating over a small circular or rectangular area is presented. The method employs eigenfunctions of angular regions together with a least squares procedure to establish continuity. Numerical results are given.

The distribution of total pressure, stagnation temperature and gas velocity is determined for the subsonic region of a supersonic jet emerging from a solid propellant rocket motor into air at rest. Measurements of Pitot pressures in the supersonic flow region have shown that the flow follows the axis of an inclined nozzle within the order of accuracy of the measurements. Nozzle cone angle, and probably the expansion ratio, affect the jet distribution significantly.

The first part of this survey dealt mainly with the development of the earliest ideas on air resistance and their application to flight. That part of the survey paused at the achievement of powered flight during the early years of the twentieth century. Even then, the main problem remained as that of aerodynamics, a subject only vaguely understood by those aviation pioneers endeavouring to get off the ground. For them theoretical guidelines, as yet, could offer little help. One of the earliest of those very few analytically based results to achieve general acceptance was the statement that resistance is proportional to the product of fluid density, a characteristic body area and the square of the flow's speed. This result had appeared during the seventeenth century in Newton's work and had received some experimental confirmation. Newton's concurrent ‘rare medium’ concept, viewing fluid flow as a stream of disconnected particles, gained credence at that time since its results agreed with the above resistance proportionality. Yet this concept became mistakenly applied to airflow so as to predict plate resistances proportional to the square of the sine of the plates' incidence angles, a misleading result which caused confusion for at least the next century.

If two small fences are arranged approximately at right angles in plan view, the magnitude and direction of surface shear stress can be deduced from measurements of the pressure difference across each fence. Fence heights as small as 0.05 mm are easily achieved. The device is simpler to use than null-seeking arrangements, and is accurate even in the presence of strong pressure gradients, which are shown to have large effects on other types of surface obstacle in three-dimensional flow.

The governing equations for bending of truncated conical shells with multi-layer anisotropic construction are developed by a variational method. The shell is considered to consist of an arbitrary number of alternating soft and hard layers. It is assumed further that the n hard membrane layers are isotropic and may possess different elastic properties, while the (n-1) soft core layers are orthotropic in general and may take transverse shear only. The variations of stresses across the membranes are neglected, as are the surface-parallel stresses in the cores. These assumptions are consistent with those usually employed in single-core sandwich shells. The energy functional is formulated with the stresses considered as independent variables. The stresses are also dependent variables defined in the set of two curvilinear co-ordinates defining the surface of the shell. The stress resultants are introduced as constraint conditions utilising Lagrange multipliers. Successful definition of an elastic neutral surface ensures the uniqueness of shell constants and the equations obtained may be in a form comparable to that of classical shell theory. The system of equations is reduced to a form where Galerkin’s method can be applied directly.

Results are presented of an experimental investigation on a non-conical wing which supports an attached shock wave over a region of the leading edge near the vertex and a detached shock elsewhere. The shock detachment point is determined from planform schlieren photographs of the flow field and discrepancies are shown to exist between this and the one calculated by applying the oblique shock equations normal to the leading edge. On a physical basis, it is argued that shock detachment has to obey the two-dimensional law normal to the leading edges. From this, and from other measurements on conical wings, it is thought that the planform schlieren technique may not be particularly satisfactory for detecting shock detachment. Surface pressure distributions are presented and are explained in terms of the flow over related delta wings which are identified as a vertex delta wing and a local delta wing. The forces acting on the wing are calculated and are shown to be very close to the two-dimensional wedge values over a wide range of incidence. In particular, it is shown that this wing, compared to one which supports a fully detached shock wave, generates a higher lift/(pressure drag) ratio at a given lift coefficient.

Measurements are presented of the mean flow properties of some three-dimensional turbulent boundary layers re-developing after reattachment behind short separation bubbles yawed at 26.5° to the main stream. For these measurements, Rθ11 varied from about 550 to 1450. It was found that, where the pressure gradient parameter (ν/ρu3τ1)∂p/∂s was not greater than about 0.05, the flow in the local external streamline direction conformed well with empirical laws for fully-attached two-dimensional layers with regard to the mean velocity profiles, shape parameter relationships and skin friction laws, giving support to the usual assumption that these two-dimensional relationships may be applied to the streamwise flow in three-dimensional layers, subject to the limitation on the pressure gradient parameter. The cross-flow profiles, on the other hand, were not generally fitted well by the often-used representations of Mager and Johnston. The variations of the kinetic energy dissipation coefficient and the entrainment rate were deduced for one of the layers, both quantities being found to be higher than those predicted by empirical relationships for conventionally-developing two-dimensional layers. However, the energy dissipation is in fair agreement with that in a similarly re-developing two-dimensional flow.

Existing effective viscosity models which have been very valuable in the mean field closure method for turbulent boundary layer computation have shown certain undesirable limitations for certain realistic but general boundary layer flows. The more general flows usually involve non-negligible considerations of pressure gradients and such wall conditions as roughness, curvature and aspiration or transpiration in varying degrees of importance. The effects of these external and wall influences have, unfortunately, been underplayed by most existing effective viscosity models. The present model of the effective viscosity is developed for a general flow and has shown remarkable agreement with experimentation, without being any more complex than existing models.

The paper first discusses the analysis of a general vector field with the aid of the concept of the field line with its tangent, principal normal and binormal. Detailed applications are made to the motion of inviscid and of viscous fluids.

In an earlier paper the lower bounds for the pressure jumps across the bow shock waves of a supersonic transport were derived, it being assumed that all the shocks had coalesced into either the bow or rear shocks, but not that the shocks were at such a great distance (asymptotic) that they had the same strength. In this paper the results of the earlier work are developed so that the lower bounds for the pressure jumps across shock waves propagating through a homogeneous atmosphere are determined by considering bow and rear shock waves simultaneously.

Unmanned Combat Aerial Vehicles (UCAVs) are believed by many to be the future of aerial strike/reconnaissance capability. This belief led to the design of the UCAV 1303 by Boeing Phantom Works and the US Airforce Lab in the late 1990s. Because UCAV 1303 is expected to take on a wide range of mission roles that are risky for human pilots, it needs to be highly adaptable. Geometric morphing can provide such adaptability and allow the UCAV 1303 to optimise its physical feature mid-flight to increase the lift-to-drag ratio, manoeuvrability, cruise distance, flight control, etc. This capability is extremely beneficial since it will enable the UCAV to reconcile conflicting mission requirements (e.g. loiter and dash within the same mission). In this study, we conduct several modifications to the wing geometry of UCAV 1303 via Computational Fluid Dynamics (CFD) to analyse its aerodynamic characteristics produced by a range of different wing geometric morphs. Here we look into two specific geometric morphing wings: linear twists on one of the wings and linear twists at both wings (wash-in and washout). A baseline CFD of the UCAV 1303 without any wing morphing is validated against published wind tunnel data, before proceeding to simulate morphing wing configurations. The results show that geometric morphing wing influences the UCAV-1303 aerodynamic characteristics significantly, improving the coefficient of lift and drag, pitching moment and rolling moment.

Smart materials that are directly embedded in the rotor blade structure are an attractive concept for active blade control. A promising approach is the use of anisotropic piezoelectric strain actuators embedded in the rotor blade skin. Especially in Europe and the US this concept has been intensively investigated over the past years. A major drawback of all configurations studied so far is the high operation voltage of up to 2,000V of state of the art piezoelectric actuators. Within the Green Rotorcraft Project of the European Joint Technology Initiative Clean Sky, a new approach with a low voltage piezoelectric actuation system is investigated to demonstrate the feasibility of this technology.

A first major step in this direction was completed by conducting a centrifugal test with a model rotor blade. The objective of the centrifugal test was to demonstrate the performance of the actuation system and the structural concept under centrifugal loads by showing that the expected twist deformation can be achieved at the nominal rotation speed and different actuation frequencies.

It was demonstrated that the new actuation system is capable of operating under representative centrifugal loads. In comparison to state-of-the-art actuators (operation voltage 500V to +1500V) the new actuation system (operation voltage -20V to 120V) exhibits higher active twist performance per active area.

A summary is given of some of the more important experimental results relating to the noise radiated from a cold subsonic turbulent jet. These are then related to the predictions of Lighthill's general theory of aerodynamic noise.

The development of the gas turbine is so rapid and the thermo-dynamic ingenuity which is being lavished upon it at the present time is so imaginative and varied that the words “ in its forms ” which appear in the title to this paper can mean as much or as little as you please. Partly because I want to limit the scope of this paper to developments which might be expected to be in service within the next five years, and partly because I am frankly not sufficiently acquainted with the characteristics of many of its more advanced forms, I am going to confine myself to a discussion of the effects upon the speed and economy of commercial aviation of the two simplest and immediate variants of the gas turbine— the simple jet-producing turbine and the simple propeller-driving turbine.