Flow separation followed by aerodynamic stall limits the operation of aircraft. Expanding the flight envelope of aircraft has been a goal of aerodynamicists for decades. This work presents findings from tests in the Oregon State University wind tunnel investigating the effectiveness of a passively actuated suction-surface flap on membrane wings. Experiments were conducted on a rigid plate and membrane wings with and without a pop-up flap. All wings had an aspect ratio of 2, while membrane pre-strain and Reynolds number were varied. An increase in lift at stall was observed for all testing conditions with flap deployment. The observed average increase in maximum lift varied from 5% to 15% for different test conditions. The variation in flap effectiveness is compared to membrane pre-strain, Reynolds number, and wing camber. A quadratic relationship between modelled camber and flap effectiveness is observed, and an optimal level of membrane camber is found to maximise flap effectiveness.

Sir George Cayley (1773–1857), a remarkable 19th century scientist, was the first to formulate the principle of powered flight when in 1809 he stated in .his paper ‘On Aerial Navigation’: ‘To make a surface support a given weight by the application of power to the resistance of the air’. More specifically, Cayley must be considered the founder of the science of aerodynamics, while also being the first to design and build and successfully fly gliders, thereby inaugurating the concept of the modern fixed-wing aircraft controlled by elevator and rudder.

The Melvill Jones equation for the profile drag of a single aerofoil is adapted to the case of an aerofoil in cascade, where the static pressure may be permanently raised (compressor), or lowered (turbine). A simplified procedure for measuring the drag is then described, assuming that the total pressure wake has the form of an error curve. A table of multiplying factors is given, for compressible flow up to an outlet Mach number of 0·9. Many published measurements of cascade drag have ignored this factor, with a consequent error of up to 25 per cent.

From tests on tunnel cascades one of the sources of loss in axial compressors could be corner stall, i.e. local separation of the flow in and near the corners between the annulus walls and the convex surfaces of the blades.

Tests on a tunnel cascade at entry Mach numbers of 0·5 and 0·6 show that, at least in a tunnel, the stalling may be cured by boundary layer control. The present tests obtain control by suction through a corner slot. The tests show that removal through the slot of 1·3% of the flow decreases the total losses of the cascade by about 25%.

A more detailed examination of the losses in the tunnel cascade without suction suggests that about three-quarters of the end losses are due to the separation and reverse flow. Much of the remaining end loss appears to result from the fact that the low energy stream tubes in the entry boundary layer would experience a disproportionately large increase in cross-sectional area during diffusion, even in attached two-dimensional flow. The suction corner control can eliminate the greater part of the end losses, by eliminating the separation.

In general, an airship is equipped with hybrid-heterogeneous actuators: the aerodynamic surfaces, the vectored propellers and the buoyant ballonets. The aerodynamic surfaces have high efficiency in attitude control at high speed. However, vectored propellers are also introduced here for attitude control under the special working condition of low airspeed. Due to the lower thrust-to-weight ratio, the composite control of hybrid-heterogeneous actuators is the primary object in controller design for an airship. In composite attitude control, first the attitude moment allocation between aerodynamic control surfaces and vectored propellers is designed according to different dynamic airspeed, to achieve the smooth motion transition from low to high airspeed, then the weighted generalised inverse (WGI) is used to design the reconfigurable actuator allocation among the homogeneous multi-actuators, where the authority of every actuator can be decided by setting the corresponding value of the weight matrix, thus the control law is unchanged under different actuator configurations. Taking the mid-altitude airship as an example, the simulations of position control, trace tracking and altitude control are provided. Simulation results demonstrate that the attitude moments allocation obtains moment distribution between the aerodynamic surfaces and the vectored propellers under different airspeeds; the reconfigurable actuator allocation achieves a good distribution and reconfiguration among homogeneous actuators, thereby enhancing the reliability of the control system.

This paper describes the market for air vehicles that offer Very Short Take Off and Landing (VSTOL) capabilities, also known as VSTOL aircraft. A commercial VSTOL aircraft will only succeed if it provides profits to both manufacturers and airlines. Being able to operate out of shorter fields offers airlines greater flexibility and therefore provides a benefit to the airline. However, adding VSTOL performance also adds cost. Successful VSTOL aircraft designs must balance the added cost against the additional benefits VSTOL performance will provide.

The first problem in any new VSTOL design is determining what field length provides the maximum benefit relative to cost. In order to be able to determine this one needs to know not only what it will cost to design and build the new aircraft, but its value to the airlines, as reflected by the demand for it and the price they are willing to pay. Only then is it possible to design the aeroplane the market wants at a sustainable price and in sufficient quantities to optimise the manufacturer’s return on investment.

This paper is a comprehensive parametric analysis of helicopters and aeroplanes that shows how much the market values short field performance by determining how much airline customers have been paying for it and how the market reacts to changes in its price. The analysis simultaneously considers over 230 currently produced models of fixed wing and rotary commercial aircraft, ranging in size from a Cessna 172 to an Airbus A380 and with balanced field lengths ranging from less than 150 to over 10,500ft.

The data show that a single, statistically significant demand curve exists for all of these vehicles, with prices that range across several orders of magnitude. This curve consists of several segments or sub-markets including those for regional and rotary wing aircraft. Both of those markets offer vehicles over a wide range of prices. The analysis shows that VSTOL price is a function of the balanced field length provided by the vehicles within these market segments. Comparing demand and sustained price against cost reveals particular field lengths for which it will be possible to make a profit building VSTOL aircraft.

Because the sustained price for a VSTOL aircraft is a function of its field length performance, and the quantity of a model sold is a function of its price, it follows that it is possible to predict the profit potential for a given VSTOL design in advance of its development. A thorough study of potential configurations therefore offers manufacturers the ability to design VSTOL aircraft that provide the maximum possible profit while fully satisfying the requirements of their airline customers.

The supersonic and subsonic jets are both considered by the method used by Miles for an oscillating aerofoil between wind tunnel walls.

The present problem and that of Miles are special cases of the flow past an oscillating aerofoil between porous walls, and provide a useful check for results in the porous wall problem. It may be assumed that the ratio of the change in pressure across a porous wall to the normal velocity at the wall defines a quantity called the porosity parameter, which is a constant at any particular point on the wall. For the special cases of the porosity parameter being everywhere infinite and zero, the porous wall corresponds to the solid wall and the free jet boundary respectively.

The solution of the general heat transfer problem for typical aircraft-type structures usually involves a finite element approach involving the solution of a large number of simultaneous equations. This may be performed by calculation using a digital computer or with an analogue simulator. For calculations involving both spatial and time intervals considerable care must be exercised to obtain solutions which are not divergent or oscillatory and this usually involves very small time increments and a large total computation time.

Aerial top-dressing, or the spreading of inorganic fertilisers from aircraft, introduces stringent flying conditions at extremely low altitudes. During release of the fertiliser there are large load and centre-of-gravity changes, and aerodynamic effects the cause of which has hitherto been conjectural. The experiments described in this paper were aimed at obtaining a knowledge of the aerodynamic effects of the cloud of falling material beneath the plane, and the effects of air flow through the empty hopper. The implications upon aircraft design are discussed. The high accident rate experienced in this industry emphasises the necessity for the production of a suitable aircraft designed solely for top-dressing.

This lecture does not set out any new method or offer any ideas to enable the design office to start cutting down the weights of new projects.

The last lecture dealing with the weight of aircraft was given by Major T. M. Barlow, M.Sc, M.Inst.C.E., M.I.Mech.E., F.R.Ae.S., in 1929.

Simultaneous equations of three or more variables are notoriously trouble some to solve numerically and, moreover, are frequently sensitive for certain relative values of the variables. This usually precludes the use of the slide rule and in consequence resort has to be had to accurate working to an increasing number of decimal places involving much labour even with a calculating machine.

A method is here given which is applicable to certain classes of equations of frequent occurrence in mathematical physics and engineering science, which method is a combination of what is knovwi as the Iteration process and the means of applying this principle to end moment distribution devised by Professor Hardy Cross.

The service life of aircraft structure includes the fatigue life and calendar life. The Aircraft Structural Life Envelop (ASLE) is a safe and reliable life scope of aircraft structures in service. The specific steps to establish the ASLE are developed, and a residual life prediction method for aircraft structure under service environments is established by combining the ASLE with the Miner theory. Furthermore, a service life extension method of aircraft structure is proposed based on a scope extension of the ASLE, including methods based on reliability analysis and structural repair. Finally, an application example of the ASLE is presented.

In recent years, aerodynamics and thermodynamics have found common ground in the specialised field of gas dynamics. Developments in this subject have led to a much more complete and widespread knowledge of subsonic, sonic and supersonic flow of gases in the conventional type of convergent-divergent nozzle. When the back pressure is raised above the value against which the nozzle is designed to discharge, oblique and then normal shock waves are set up in the divergent cone at a position along the nozzle axis determined by the magnitude of that back pressure. The gas which has crossed the shock wave is subjected to a process of subsonic compression.

In this paper a theoretical investigation is made of the changes in pressure, temperature, density and Mach number which occur across a normal shock wave, when the position of the wave varies along the nozzle axis. The investigation illustrates the effect of change of medium, for which the relevant property is the ratio of specific heats. This ratio for certain polyatomic gases may approach unity (e.g. for Dichlorodifluoromethane CCl2F2, in gaseous form, γ=1·06), and for the inert monatomic gases γ= 1·667. The analysis is made non-dimensional by expressing such quantities as gas pressure, temperature, and density at any given position along the nozzle axis relative to the values of the particular parameter at entry to the convergent section.

The twenty-first Louis Blériot Lecture was given in Paris on 5th April 1968 under the auspices of L’Association Françaises des Ingénieurs et Techniciens de L'Aéronautique et L'Espace. The lecture “Some Thoughts about the Future of European Aeronautics” by Handel Davies, CB, MSc, CEng, FRAeS, FAIAA, was read on his behalf in French by General Lecamus, FRAeS. The Lecture was attended by the President, Mr. M. B. Morgan, CB, MA, FRAeS, the Secretary of the Society, Dr. A. M. Ballantyne, OBE, TD, PhD, by several members of Council and a large and distinguished French audience.

The Chair was taken by Monsieur Robert Blum, President of AFITAE. Following the lecture a dinner was given by AFITAE.

The Thirty-Sixth Wilbur Wright Memorial Lecture was delivered before the Society by Mr. A. Gouge on Thursday, 27th May, 1948 at 6 p.m. at the Institution of Civil Engineers, Great George Street, London, S.W.I. The chair was taken by Dr. H. Roxbee Cox, President of the Society.

Pilots are sometimes not provided with sufficient information to avoid go-arounds or other operational disruptions that result from low-level wind disturbances. We identified issues with existing windshear alerting systems and developed three types of airport low-level wind information systems to enhance pilot situational awareness of wind conditions by providing landing aircraft with quantitative and visualised wind information for ultimately mitigating air service disruptions due to low-level wind disturbances. The three systems, Airport Low-level Wind Information (ALWIN) and Low-level Turbulence Advisory System (LOTAS), both of which use Doppler radar/lidar, and Sodar-based Low-level Wind Information (SOLWIN), which uses Doppler SOnic Detection And Ranging (SODAR), have different costs and capabilities that allow the most cost-effective system to be selected for an airport according to its scale and local weather characteristics. This paper presents the operational concepts of our newly developed airport low-level wind information systems and describes their validation.

A two-dimensional vortex particle model is used for studying the development of the vortex wake generated by the wing of an aircraft in the Trefftz plane. Two-dimensional, finite-area vortex structures simulate the initial vorticity distribution at a near-field cross-section of the wing, as provided by either measurements or simulations. The code is used for studying the effect of weak or strong counter-rotating vortices on the development of an aircraft wake. Application in a three-vortex configuration, consisting of the tip and flap vortices, plus a weak negative vorticity sheet lying between them, supports the hypothesis of Graham that the merging of the tip and flap vortices is prevented, because in such an arrangement the counter-rotating vorticity sheet is wrapped around the primary vortices. The present results indicate that even in the far field the tip and flap vortices remain distinct and follow a helical trajectory with large-pitch and small-radius. The code is next applied to the two-vortex system of Savas’s triangular wing, in which the circulation of the flap vortex is comparable to that of the tip. Although a 2D analysis is inappropriate for stability analysis it is still useful for a quick qualitative investigation. Results indicate that the flap and tip vortices follow a helical trajectory with large-pitch but also very large radius. During one period, the flap vortex covers a span wise distance equal to the wing span. Such a flow has not been observed in flight or in laboratory tests. Actually, Ortega et al, who studied experimentally the triangular wing of O. Savas, found out that before concluding one full spiral, the vortices are literally destroyed (rapid spreading of their vorticity) by an instability mechanism.