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The purpose of this paper is to rise the autonomous flight performance of the small unmanned aerial vehicle (UAV) using simultaneous tailplane of UAV and autopilot system design.

A small UAV is remanufactured in the UAV laboratory. Its tailplane can be changed before the flight. Autopilot parameters and some parameters of tailplane are instantaneously designed to maximize autonomous flight performance using a stochastic optimization method. Results found are applied for simulations.

Benefitting simultaneous tailplane of UAV and autopilot system design process, autonomous flight performance is maximized.

Authorization of Directorate General of Civil Aviation in Turkey is required for UAV flights.

Simultaneous tailplane and autopilot system design process is so useful for refining UAV autonomous flight performance.

Simultaneous tailplane and autopilot system design process fulfills confidence, high autonomous performance, and easy service demands of UAV users. By that way, UAV users will be able to use better UAVs.

Creating a novel technique to recover autonomous flight performance (e.g. less overshoot, less settling time and less rise time during trajectory tracking) of UAV and developing a novel procedure performing simultaneous tailplane of UAV and autopilot system design idea.

IN the May, 1933, number of AIRCRAFT ENGINEERING there was published an article by Korvin‐Kroukovsky on the uncontrolled tail spin, in which one of the main points put forward was that during a spin both the body and the tail were very largely shielded by the wings. It was then argued from this that danger might be expected more especially in the low‐wing monoplane type and also the biplane with big forward stagger of wings. At the time there was very little direct evidence on the effect of wings on body and tail moment although an early report on the “ Bantam,” an unstaggered biplane which had spun into the ground, had shown pretty clearly that it was the tailplane which was responsible for the complete blanketing or even reversal of the fin and rudder in a flat spin, and that this could be remedied by raising the tailplanc.

A Discussion concerning the Use of Wind Tunnel Results and Flight Test Measurements in the Prediction of Aerodynamic Loads for Stressing Purposes in the Aerodynamics Department of the Weybridge Division of British Aircraft Corporation. The responsibility for the prediction and issue of aerodynamic loads for stressing purposes at the Weybridge Division of British Aircraft Corporation is carried by the Aerodynamics Department. The arguments for and against this arrangement are briefly examined. One of the main arguments in favour is the facility with which wind tunnel tests can be instigated and controlled. The use of wind tunnel tests specifically designed to give aerodynamic loading data and their relation to estimation using theoretical and semi‐empirical methods is fully discussed and illustrated. The confirmation of design estimates by full scale in‐flight load measurement is described and the usefulness of in‐flight measurements as a design tool on subsequent aircraft of a similar type is discussed.

THE B.206 had its origin in the B.206X experimental aircraft which was built in 1961 to obtain aerodynamic and structural data for the production series, and it first flew in August 1961.

THE mathematical theory of longitudinal stability appears to have been adequate to explain the salient features of the behaviour of aeroplanes in longitudinal motion. In general the provision of a stable slope to the static pitching moment curve has been found in practice to fulfil all requirements, and although increasing oscillations do on occasion occur, they are on the whole surprisingly rare. The reasons for this are fairly well recognised and are briefly indicated in what follows. There is little doubt, however, that the designers' principal difficulties centre round the complex interferences between the wings and the tailplane, particularly with the air‐screw running. The downwash from the centre section in many machines, even with no engine on, is quite unpredictable in the present state of knowledge, and the calculation of the downwash due to the slipstream has not yet been successfully made even in the simplest cases. Some attempt is here made to summarise the present position.

This study aims at presenting one of the most important problems of aeroplane design, Static Longitudinal Stability, in a possibly complete but easy way.

This article deals with some of the stability, control and handling problems that have arisen as a result of drastic changes in aircraft configuration coupled with the advent of supersonic flight at high altitude. The article will be published in two parts. The present part contains a brief introduction to the subject of aircraft stability and control in addition to a description of the longitudinal characteristics of supersonic aircraft. The second part will be published in our next issue, and will deal with the lateral characteristics of supersonic aircraft. Some of the problems encountered in the design of the flying control system for this type of aircraft and an indication of the methods and techniques used for solving the various stability problems are also presented in the second part.

THE structural design of the VC10 has been based on well‐established principles and the philosophy of low stress levels in the skins and secondary load paths to meet ‘fail‐safe’ conditions. In general a multiple load path type of structure has been employed. In any one of these load paths a crack may be allowed to develop without fear of catastrophic failure. The operator will, however, be required to maintain a high standard of inspection of the structure in order to prevent cracks developing concurrently in these multiple load paths. The operating economy will be greatly enhanced since major components will not require to be changed until the whole of their useful life has been utilized. This was not the case in structures of the ‘safe‐life’ type owing to the large factors which had to be imposed on life.

Extreme loads are generated in aircraft flight manoeuvres. Among different manoeuvres considered in this paper are motions following a sudden deflection of elevator and response to a vertical gust. Airplane was assumed to be a rigid body of three degrees of freedom in symmetrical motion. Elevator deflection was either of the step change type, or of the sinusoidal type, gust was assumed to be either of the step change type or harmonic, with a gust cycle time corresponding to the time to travel a distance equal to 25 Mean Aerodynamic Chord. In all cases a jump type elevator deflection was assumed to last for 1 second, whilst the airplane response was observed for 3 seconds. The airplane motion, its velocities, accelerations and load acting on the tailplane were calculated by means of numerical integration of the ordinary differential /of motion.

A detailed account of the wing and flap design, tailplane considerations and the adoption of speed brakes. FOKKER has already demonstrated with the F.27 Friendship the economical advantages of an aerodynamically clean aircraft. The economics of the Fellowship is still more sensitive to drag because of the much higher cruising speed and the fact that it is jet propelled.