Wind tunnel data have been obtained from a flapped tanker wing and receiver aircraft model at varying vertical separation and the results compared with theory. In the aerodynamic model, the tanker wing is represented by a pair of horseshoe vortices from the wing and flap tips and an allowance is made for the self-induced vertical displacement of the trailing vortices at the receiver aircraft position. The aerodynamic loads on the receiver are determined by the vortex-lattice method and lifting-line theory, although an approximate method is used to determine the side force on the fin. Data were obtained from open and closed test sections in order to estimate the wind tunnel boundary interference effect. In the longitudinal case, significant differences were obtained between theory and experiment. Two reasons for the differences are the assumption that the trailing vortices are fully rolled up and the neglect of viscous decay of the vortices. The lateral aerodynamic interference was determined by banking the tanker wing and displacing it sideways and by yawing the receiver aircraft model. In the case of the rolling moment due to sideways displacement, which is the most significant lateral aerodynamic interference term, the wind tunnel boundary interference is highly significant and the difference between theory and experiment is large due to incomplete roll up of the trailing vortices. The theoretical and experimental trends are similar although the theory overpredicts the rolling moment due to bank and sideways displacement while the corresponding side force and yawing moment are underpredicted. As in previous tests using an unflapped tanker wing, the effect of the sidewash due to the tanker wing on the receiver in yaw is found to be relatively insignificant.

An analysis is presented of the output of a pressure probe, comprised essentially of four yawmeters, which was designed to yield vorticity information.

It is shown that if the probe is utilised in flows which have vorticity dominated by variation of either the direction or the magnitude of the velocity then the probe output may be easily related to vorticity. It is argued that such flows should be readily recognised, and they include many situations of interest involving vortex flows, boundary layers and wakes.

Earlier work had already demonstrated that the probe is useful in exploring regions of predominantly streamwise vorticity. The present work shows a result with the probe used in a boundary layer flow where the vorticity is nearly normal to the local flow.

Use of the probe in unsteady flow is briefly examined.

A considerable amount of theoretical investigation has been undertaken recently, into the stress concentrations arising in the vicinity of cut-outs in aircraft pressure cabins. Papers which analyse the stress concentrations around circular, elliptical, and square reinforced cut-outs, under various loading conditions have been presented by Mansfield, Hicks and Wittrick.

In all of these investigations, the problem of the cut-out in a cylinder has been reduced to a problem of an infinite plane sheet containing a hole, and the effects of the curvature have been generally assumed to be insignificant.

It has almost become a tradition in recent years to begin a paper on boundary layer by paying a tribute to the great Prandtl and his famous paper in which he introduced the conception of boundary layer into Fluid Dynamics.

The year 1954 is, however, not only memorable to mark the passing of fifty years since Prandtl's classical demonstration of the effect of boundary layer suction on the flow pattern around a cylinder; it is also memorable because in 1954 most convincing demonstrations of practical applications of boundary layer control for aircraft have taken place. I am referring, in particular, to the demonstrations with the Attinello flap in the United States which mark the introduction of one form of boundary layer control as an engineering and practical reality.

The flying wing supersonic transport is simply a straight wing of high aspect ratio and of sufficient size to contain passengers comfortably. Such a wing can be steered to different sweep angles and thus adapted to efficient flight over a range of Mach numbers.

The writer demonstrated such a wing by flying models at the First International Congress in the Aeronautical Sciences. The potential advantages of such a flying wing were discussed by G. H. Lee of Handley Page Aircraft at a subsequent Congress. The present paper gives recent studies of flight stability, efficiency and operational problems. It is suggested that by limiting the Mach number to twice current jet speed fuel efficiency approaching that of present day subsonic transports could be achieved. This goal would probably also require the development of variable bypass engines.

Several critical load cases during the aircraft design process result from atmospheric turbulence. Thus, rapidly performable and highly accurate dynamic response simulations are required to analyse a wide range of parameters. A method is proposed to predict dynamic loads on an elastically trimmed, large civil aircraft using computational fluid dynamics in conjunction with model reduction. A small-sized modal basis is computed by sampling the aerodynamic response at discrete frequencies and applying proper orthogonal decomposition. The linear operator of the Reynolds-averaged Navier-Stokes equations plus turbulence model is then projected onto the subspace spanned by this basis. The resulting reduced system is solved at an arbitrary number of frequencies to analyse responses to 1-cos gusts very efficiently. Lift coefficient and surface pressure distribution are compared with full-order, non-linear, unsteady time-marching simulations to verify the method. Overall, the reduced-order model predicts highly accurate global coefficients and surface loads at a fraction of the computational cost, which is an important step towards the aircraft loads process relying on computational fluid dynamics.

The paper presents a method for constructing a set of vertical flight path segments, that would compose an aircraft's vertical flight envelope, by using an aircraft performance model. This method is intended to be used for aircraft flight plan prediction and optimisation algorithms. The goal is to reduce the volume of recurring segment performance computations currently required for flight plan prediction or optimisation. The method presented in this paper applies to a free-flight scenario. The flight-path segments composing the vertical flight envelope belong to one of the unrestricted climb, constant-speed level flight, step-climb and continuous descent segments, performed at the consigned climb, cruise and descent speed schedules and at the consigned air temperature values. The method employs an aircraft model using linear interpolation tables. Nine test scenarios were utilised to assess the performances of the resulting flight envelopes as a function of the number of cruise altitudes and descent flight paths. The set of evaluated performance parameters includes the range of total flight times and still-air flight distances, and the vertical profiles describing the minimum and maximum flight times, and still-air flight distances. The advantages of the proposed method are multiple. First, it eliminates the need for repetitive aircraft performance computations of identical vertical flight plan segments, and provides the means for quick retrieval of the corresponding performance data for use in the construction of a full flight plan. Second, the vertical flight path look-up structure and the vertical flight-path graph describe a set of vertical flight paths that consider an aircraft's and flight plan's configuration parameters and cover its maximum flight envelope. Third, the look-up structure and the graph provide the means for rapid and clear identification of the available options for constructing a flight-plan segment, as well as for detecting the points associated with changes in the flight phases, including climb, cruise, step-climb and descent.

This paper describes a detailed experimental investigation of turbulent heat transfer in a radially rotating tube with particular reference to its application to the design of cooled turbine rotor blades. The paper focuses on the strategy adopted, the description of the apparatus used, the method of data processing and a selection of measurements which illustrate the manner by which Coriolis force and centripetal buoyancy force interactively affect the local heat transfer along the leading and trailing edges of the tube.

As well as re-confirming the fact that Coriolis force and centripetal buoyancy have a significant influence on the forced convection mechanism present due to the through flow, a number of new experimentally-based observations are presented. An empirical correlation, which is physically consistent, has been developed which permits the interactive effect of Coriolis force and centripetal buoyancy on forced convection to be evaluated and quantified.

The work has been motivated by the need to understand the general effect of rotation on the performance of the internal cooling airways used in gas turbine rotor blades.

In this paper the shock waves of a slender aircraft are considered at a great distance from the aircraft so that all the shocks have coalesced into either the bow or rear shocks, but not at such a great distance (asymptotic) that the two shocks are of equal strength. Lower bounds are determined for the pressure jump across the bow shock subject to various aircraft constraints, and the pressure jumps across both the bow and rear shocks are determined at off-design conditions. The work described is considered to be a preliminary investigation and the results are such that further work seems to be warranted.

For a simple undamped vibrator driven by a force whose frequency varies linearly with time, simple algebraic expressions are derived which give values for the maximum amplitude performed by the vibrator and for the frequency of the driving force where the maximum displacement occurs. Driving forces of both increasing and decreasing frequencies are considered.

The Swiss Airforce is operating F/A-18C/D Aircraft since 1997. Since the aircraft’s structural design is different from the version operated by the US Navy it was necessary to carry out a structural integrity study (ASIP) which was done by The Boeing Company in St. Louis. To validate this study a full scale fatigue test facility was build at RUAG and operated from 2003 to 2005. When operating this facility difficulties were encountered with the aerodynamic loads data provided by Boeing (insufficient, not well documented, questionable data). As a result RUAG looked for alternative methods to provide the aerodynamic loads, and a large investment was made in the development of a Computational Fluid Dynamics (CFD) tool. The Navier Stokes Multi Block (NSMB) solver, which was developed in an international collaboration, was adopted. In a first phase the code was validated by comparing results of CFD calculations with wind-tunnel results, results from literature and flight test data results. In the second phase, discussed in this paper, a Fluid Structure Interaction (FSI) tool was developed to permit unsteady aero-elastic simulations. Particular attention is focused on the vertical tail since this component of the F/A-18 fighter is very sensitive to fatigue due to unsteady loads generated by buffeting phenomena.

A study of the laminar boundary layer on a yawed infinite wing in a compressible fluid is made. A method is presented for the investigation of compressibility effects by direct solution of the linearised equations of motion, first considerably simplified by the extension of a transformation due to Illingworth and Stewartson, which assumes a heat-insulated surface and a fluid of unit Prandtl number and a coefficient of viscosity which is proportional to the absolute temperature. As an example of this procedure the boundary layer near a stagnation point is calculated. The simplifications necessary for an extension of the momentum method, which is available for a yawed infinite wing in an incompressible fluid, are discussed, and this method is recommended for a first evaluation of the effect of compressibility.

When I received the invitation to deliver this Memorial Lecture, and “Turbulence” was chosen as the subject, I remembered that my good friend, G. I. Taylor, gave a very stimulating talk on the same subject ten years ago. It was his G. I. Symons Memorial Lecture before the Royal Meteorological Society. He succeeded in fulfilling his job without using one mathematical symbol, the President of the Society, Sir Gilbert Walker, insisting very definitely that “turbulence without tears” was what he wanted. I do not think that I can duplicate Taylor's achievement completely; however, I shall try to reduce the mathematics to a minimum.

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.

Aerodynamic performance degradation has been considered a critical hazard to aircraft due to flying in heavy rain. In this work, a two-way momentum coupled Eulerian-Lagrangian approach is developed to study the aerodynamic performance of a two-dimensional (2D) transport-type NACA 64-210 cruise and landing configuration aerofoil as well as a three-dimensional (3D) NACA 64-210 cruise configuration rectangular wing in heavy rain environment. Raindrop impacts, splashback and formed water film are modeled. The steady-state incompressible air flow field and the raindrop trajectory are calculated alternately by incorporating an interphase momentum coupling term through a curvilinear body-fitted grid surrounding the aerofoil/wing. Our simulation results agree well with the experimental results and show significant aerodynamic penalties for all the test cases in heavy rain. Rain-induced premature boundary-layer transition and separation are observed and details of the raindrop splashback effect on the boundary layer are examined. A 1° rain-induced decrease in stall angle-of-attack is predicted. An uneven water film upon the wing surface is observed and its effect on the wing surface roughness is also examined.

A new integral equation, and the corresponding boundary-element method, is presented for the aerodynamic analysis of compressible, subsonic, potential flows around bodies in arbitrary motion, with applications to helicopter rotors in hover. The formulation is closely related to that introduced by Morino for the aerodynamic analysis of aircraft: in that paper, the surface of the aircraft is assumed to be in arbitrary motion with respect to a frame of reference that moves in uniform translation. Such a formulation is very useful for the analysis of aircraft undergoing small oscillations with respect to a reference configuration. However, for the analysis of more complex motions (e.g. helicopter rotors in forward flight), it is more convenient to use a frame of reference that moves in arbitrary motion (in particular, for the case of a body moving in rigid-body motion, a frame of reference rigidly connected with the body).

An analysis technique capable of simulating the effect of engine failure during takeoff from offshore platforms is presented. Use is made of inverse simulation whereby the control actions required for a subject helicopter to follow a particular trajectory can be established. A mathematical representation of the towering takeoff procedure, and details of the modified inverse simulation technique needed to cope with the modelling of an engine failure are described. Detailed piloting accounts of the strategy used to fly the towering takeoff (with and without engine failures) are also given and are used to give qualitative validation of the analytical approach. Simulation results for a single engine failure of a transport helicopter during critical phases of a towering takeoff are presented. Finally, some directions for future work and potential applications of the technique are discussed.

Hypersonic conical flows over delta wings are treated in the thin-shock-layer approximation due to Messiter. The equations are hyperbolic throughout, even in regions where the full equations are elliptic, and have not hitherto been solved for flows with attached shock waves. The concept of the simple wave has been used to construct a class of solutions for such flows; they contain discontinuities in flow variables and shock slope but, for the case of flow over a delta wing with lateral symmetry, agreement with results of numerical solutions of the full equations is good. The method is applied to plane delta wings at yaw, and to wings with anhedral and dihedral. For the flow at the tip of a rectangular wing, it is shown that two distinct solutions may be constructed.

The first centennial of man's first powered flight, which was performed by Wilbur (1867–1912) and Orville (1871–1948) Wright (Fig. 1) will be celebrated in 2003. These brothers' unique and trendsetting enterprise, their skills to develop, build and commercialise controllable and load-carrying flying machines, formed the first example of a private interdisciplinary aerospace development and design initiative. It was at the same time a rare example of entrepreneurial engineering. Their work involved wind-tunnel testing of lifting devices and full-scale tests of major structural components. Equally if not more importantly they developed an essential propulsion system consisting of a lightweight 36hp engine with four cylinders in line and a fuel injected carburettor. This engine propelled two pusher propellers through a drive system of chains. Moreover they developed the systems to control their flight. The twistable wing tips that control rolling and heading during the flight were the last remains of the former trend to mimic bird flight.

This paper presents a novel method for quantifying the effect of ambient, environmental and operating conditions on the progression of degradation in aircraft gas turbines based on the measured engine and environmental parameters. The proposed equivalent operating time (EOT) model considers the degradation modes of fouling, erosion, and blade-tip wear due to creep strain, and expresses the actual degradation rate over the engine clock time relative to a pre-defined reference condition. In this work, the effects of changing environmental and engine operating conditions on the EOT for the core engine booster compressor and the high-pressure turbine were assessed by performance simulation with an engine model. The application to a single and multiple flight scenarios showed that, compared to the actual engine clock time, the EOT provides a clear description of component degradation, prediction of remaining useful life, and sufficient margin for maintenance action to be planned and performed before functional failure.