After a brief discussion of alternative ways in which the drag of an aircraft wing can be derived theoretically, attention is focused on a technique whereby the separate ‘far-field’ components of drag — viscous, trailing-vortex (induced) and wave — are calculated separately. In particular, a new approximate method is described for estimating the wave drag. Based on an exact two-dimensional analysis involving the flow conditions just upstream of the shock wave, a simple formula is derived which, to first order, involves only a knowledge of the Mach number distribution on the surface of the aerofoil ahead of the shock and of the surface geometry at its foot. The accuracy of this formula is assessed for aerofoils by comparison with more exact theoretical results and with experiment. A ‘strip theory’ extension to swept wings is proposed and illustrated by applying it to a particular transport-type wing body combination. Using experimental pressure measurements as input, all three components of drag are estimated theoretically, and by adding their sum to separate balance measurements of the body drag comparisons between ‘theory’ and experiment for the overall drag can be made. These show a satisfactory standard of accuracy, the error varying between —5% and — 1% of the total drag over a wide range of Mach number and lift coefficient.

Two hundred years ago the first manned balloon launch was attempted. On 21st November 1783 the hot air balloon, Montgolfiere, carried two men, the adventurer Pilatre von Roxier and the Marquis D’Arlandes, nine kilometres during their twenty minute flight.

The history of flight continued with many ambitious and inventive men. Among them were Graf Zeppelin, famous for airships, Otto Lilienthal for motorless flight, and the Wright brothers for the first powered flight. Since then the development went further until today large scale aircraft have become commonplace. Tomorrow, space flight could be just as common. Who dares say no to the technology that has brought us today’s living comforts?

The following contribution deals with an actual problem in present day hot air balloon construction.

The variation of the Crack Stress Intensity Factor (SIF), for two cracks growing from one hole only in an array of holes in an isotropic material under uni-axial tension has been studied. The effect of the fastener hole ahead of the growing crack was found to give an increase in the SIF only when the crack was within about one and a half diameters of the approached hole edge, compared with the level predicted by considering a single central hole in a plate. An equation suitable for inclusion in a crack growth program has been derived from the research results.

This paper discusses the cost of high technology and the benefits to commercial air transport. The manufacturer's dilemma today is that whereas one could pay a 16% price premium to obtain a 10% reduction in fuel cost in the late 1970s, in 1985 that price premium has diminished to only 2½/2% to achieve a 10% fuel saving. Consequently, we must obtain the benefits of high technology at an extremely low cost.

Four areas are discussed in some depth: Aerodynamics, including the effects of computational fluid dynamics, wing design, and laminar flow. Structures, including the effects of new materials, both metallic and non-metallic, and the automated tools necessary to build the non-metallic structure. Avionics, including both fly-by-wire and fly-by-light, and the improvements due to high speed digital computers as well as flat panel flight deck displays. Propulsion systems, including the effects of high bypass ratio, high pressure ratio turbofan engines, and single and counterrotation turbopropellers. The final result is an integration of these four areas of technology into a composition transport aircraft and the resulting performance.

We all tend to think of the aviation industry as exciting high technology and eagerly look forward to the next new developments. Unfortunately, the air transport industry has just been through a most agonising depression on a worldwide basis, and their primary thoughts are, 'How can we keep those manufacturers with their new toys at bay until we can recover?' Certainly the manufacturers respect this view because we, for one, have just spent several billions to roll out three new products, and high technology has been expensive. To illustrate, if we compare the cost elements of a 727 of 1970 with one of 1980 (Fig. 1), we see the 68% increase in fuel cost (Fig. 2) that certainly drove the designs of the 757 and 767 airplanes. But if we look further at the cost element breakdown for a 1985 757, we see the large reduction in cash cost due to the high efficiency of its Rolls-Royce engine and the technology of its new wing, structures, and systems. What we often fail to perceive properly is the very large change due to the ownership cost of this high technology aircraft.

This paper deals with control of V/STOL aircraft in powered-lift flight. A review of representative operational mission tasks is presented for the take-off, transition, hover and landing flight phases. Characteristics inherent in the basic airframe and propulsion system and the behaviour associated with control augmentation systems are considered. Demands for augmented stability and control response to meet certain mission operational requirements are discussed. Experience from ground-based simulation and flight experiments that illustrates the impact of augmented stability and control on aircraft design is related by example.

A linearised two-dimensional incompressible potential flow theory for two-element aerofoil sections is developed. It is capable of predicting the effects of angle of attack, flap deflection, camber, thickness, stagger and overlap of aerofoil elements on the forces. These effects are summarised in integrals which are analogous to the one-element thin aerofoil theory Munk integrals. Analytical expressions for the forces on tandem NACA 23012 aerofoils have been derived and results are presented. Comparisons are also made with a realistic slotted flap configuration with overlap for which experimental data is available. The linearised theory is seen to correlate with these real flow results as well as and better than it does in the one-element regime.

Measurements of the turbulent boundary layer on a sidewall of the RAE 8 ft × 8 ft wind tunnel are described. Velocity profiles and wall shear-stress measurements were obtained for a Mach number of 0·8 over a Reynolds number range of over a decade up to a value of nearly 400 million.

A skin-friction relationship is deduced together with a description of the velocity profile and boundary-layer growth based upon the velocity-defect law using kinematic parameters for constant-pressure flows without heat transfer.

Previous data for Mach numbers up to 2·8 obtained in the same tunnel are shown to be well matched with the present analysis and the use of simple scaling factors for compressible flows is confirmed.

The results of an experimental evaluation of three assessment criteria for boundary layer two-dimensionality are presented. The criteria include the so-called ‘two pin test', the momentum balance principle and the cross-stream variation in local skin friction coefficient. Extensive spanwise measurements in the flows investigated highlight the relative merits and limitations of the criteria imposed.

The momentum balance method is found to be the most informative technique, with the proviso that the test data must be interpreted with some care if measurements are restricted along a single streamwise axis.

The two pin test is found to provide a worthwhile complimentary addition to momentum balance data, whereas the spanwise variation in local skin friction coefficient is less conclusive as regards flow two-dimensionality.

Excepting the case of adverse pressure gradient flows, the momentum balance equation is seen as a viable experimental means of estimating the local skin friction coefficient in boundary layers where measurement is problematical, as in the case of transitional boundary layers. In such cases, the two pin test could be used as the sole means of confirming the flow two-dimensionality.

Hardware and software limiting of the velocity and displacement of motion platform hydraulic jacks is an important source of spurious cue generation in the important frequency range below 2 Hz. Until recently the cross-coupling effect of such limiting has been ignored in motion-drive-software. By setting software limits and scaling jack demands with respect to this limit and the maximum jack demand, a limiting strategy which maintains the directional fidelity of the acceleration cue is achieved. This simple scheme includes a soft limiting approach to reduce motion jerks. In addition the scheme requires negligible extra computing time or storage over the traditional schemes.

This report gives extended data for the flow past roughness trips to provoke boundary-layer transition. This first Part reports work on wire trips giving data for the forces on the wire, the positions of recovery and effective origin of the turbulent layer, the length of the separation bubble and the beginning and end of transition. Parts 2 and 3 describe similar work on spherical roughness trips, used singly and in rows, looking in more detail at the spanwise variations in the downstream boundary-layer.

Following Part 1, experimental data is given for the boundary-layer flow past single roughness trips of spherical shape. Correlations are given for the start and end of transition, for the recovery position and for the effective origin of the turbulent layer. There is evidence of a spanwise array of vortices which are found to influence both the downstream turbulent layer and the wedge growth of turbulence.

An exact integral formulation of the two-dimensional full potential equation is presented. The well developed standard surface panel method based on the linear Prandtl-Glauert equation is extended by means of sources in the field surrounding the configuration, allowing the calculation of transonic flows.

A major attraction of the Field Panel Method is that solutions can be obtained using standard surface panelling together with a field grid which has no special form. The surface panels are used to satisfy the boundary conditions so that the field grid does not have to be surface-fitted which is a requirement for existing field methods. Further, the grid need only be in a region close to the configuration where nonlinear compressibility effects are non-negligible and need not extend to the far field as generally required. Grid generation is therefore trivial allowing solutions for arbitrary configurations. Convergence, especially for flows with shock waves, is enhanced by use of a modified Approximate Factorisation scheme. Results for single aerofoils are compared with Garabedian-Korn solutions and show good agreement; and to demonstrate the ease of grid generation, results for multicomponent aerofoils are also presented. Finally, an outline is given of how the method can be extended to threedimensions.

The work was carried out in conjunction with the Department of Applied Mathematics, University of Leeds together with Procurement Executive, Ministry of Defence.

Details of a visual method for indicating boundary layer transition are given, with particular reference to tests at high Reynolds number in the RAE 5 metre pressurised low-speed wind tunnel on a slender-body model. The method is compared with other techniques on the basis of ease of use and quality of the indication.

Following Parts 1 and 2 of this experimental study the effects of rows of spherical roughness elements upon transition of the boundary layer are now described. Data presented includes that for the start of the end of transition, the effect of spacing of the roughness elements, the effective origin of the turbulent boundary layer and its development and the spanwise variations across the turbulent boundary layer.

We consider the effect of atmospheric disturbances, such as wind, shears, turbulence, wakes and downflows, on aircraft performance (section 1), recalling the minimum of aerodynamics (section 2) necessary for the calculation of the effects on flight mechanics. For example, it is shown (section 3) that the relative lift change, due to wind, shear or both, coincides with the disturbance intensity G, defined as the instantaneous vertical acceleration, measured in g's, that an aircraft will experience as a consequence of atmospheric disturbances, assuming constant velocity and attitude. The disturbance intensity is a measure of the total force that the atmospheric disturbance can exert on the aircraft; this force may cause changes in normal acceleration, flight velocity and angle-of-attack, either isolated (section 5) or combined (section 6), in straight and level flight; in the case of a horizontal turn (section 7), the disturbance intensity specifies the changes in instantaneous turn rate, velocity, radius (section 8), acceleration and bank angle (section 9) due to atmospheric perturbations, and their cumulative effect over time in (section 10) deforming and displacing trajectories. These effects (section 11) are quantified by formulas and illustrated in graphics, and we indicate throughout the flight mechanical consequences of disturbances with the critical intensities G1 = — 0·17 and G2 = — 0·42; these are the disturbance intensities which, if uncompensated, would be just enough to cause an aircraft to stall (section 4) respectively at unstick on take-off and on approach to land. We conclude (section 12) with a discussion of the use of the disturbance intensity as a parameter in the reduction and analysis of flight data.

Measurements are presented of the frequency of vortex shedding from two circular cylinders in a staggered arrangement. Seven different cylinders were used in turn as the upstream cylinder and the ratio of the diameter of the upstream cylinder to that of the downstream cylinder ranged from 0·00422 to 0·571. For a constant free-stream velocity changing the upstream cylinder changed the frequency of vortex shedding from the cylinder. It was found that the frequency of vortex shedding from the downstream cylinder N synchronised with that of the upstream cylinder n over a range of the latter frequency. When the frequencies were synchronised the ratio of the frequencies was given by n/N = 2. This behaviour is characteristic of a lock-in process.

Lock-in was found at each of the four upstream cylinder positions tried. However, it was only when the upstream cylinder position was such that the upstream cylinder only weakly affected the flow about the downstream cylinder that lock-in was accompanied by a decrease in the base pressure. The time-mean drag force then increased with lock-in, but by a small amount only. On the other hand the fluctuating lift force increased substantially.

A method to calculate two-dimensional or axisymmetric, compressible, wall-bounded viscous flows is described. The solution procedure involves a finite-difference discretisation of the governing flow equations. A number of novel normal co-ordinate transformations are used to enable efficient use to be made of the computational grid points. The implementation of zero-, one- and two-equation turbulence models is described. Results are presented for a range of compressible boundary-layer flows, and for an incompressible wake/ boundary layer mixing flow.

This paper describes a simple method for smooth curve fitting or function interpolation. Where possible, the method uses continuous piece-wise conies in preference to cubics so as to prevent the occurrence of extraneous inflection points. The method is ideal for single-valued functions with sharply changing gradients such as occur in pressure distributions, containing shock waves. A version that is valid for multivalued functions is also described.

It is useful to have available simple prediction formulae for preliminary design work. In the case of landplane take-off such formulae exist, being expressed only in terms of a few major parameters. The flying boat situation is more complex due to the much greater and more variable total drag. By making appropriate assumptions an equivalent formula for flying boat take-off has been derived. Comparison of typical cases of the landplane and flying boat shows that whereas the take-off distance of the former is approximately inversely proportional to thrust/weight ratio to the power of 1·35, in the case of a flying boat it is inversely proportional to the square of thrust/weight ratio.

Large deflection analysis of a cantilever beam under a tip concentrated rotational load governed by a second order non-linear differential equation is solved using a fourth-order Runge-Kutta integration scheme. Initially the two point boundary value problem is converted to an initial value problem by estimating one of the two required initial conditions in an iterative process, so as to satisfy the other boundary condition. The details of load deflection curves are presented.