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WE concluded Part II of this series with the remark that a different outlook is needed for problems of control surface flutter than for those of wing flutter. There are two reasons for this. Wing flutter must be investigated carefully early on in the design of an aircraft so as to provide a safe aircraft without a severe weight penalty, whereas the weight penalty of avoiding control surface flutter is usually small, although not negligible, and modifications can often be made at short notice, so it is important to make a full investigation as late as possible before flight when all the data are available in a reliable form. The second reason is that with wing flutter, as with aileron reversal and divergence, it is usual to think of safety margins in terms of forward speed or possibly wing torsional stiffness; with control surface flutter, on the other hand, quite different types of safety factor become the rule.

The purpose of this study is to develop an active controller of both leading-edge (LE) and trailing-edge (TE) control surfaces for an unmanned air vehicle (UAV) with a composite morphing wing.

Instead of conventional hinged control surfaces, both LE and TE seamless control surfaces were integrated with the wing. Based on the longitudinal state space equation, the root locus plot of the morphing wing aircraft, with a stability augmented system, was constructed. Using the pole placement, the feedback gain matrix for an active control was obtained.

The aerodynamic benefits of a morphing wing section are compared with a wing of a rigid control surface. However, the 3D morphing wing with a large sweptback angle produces a washout negative aeroelastic effect, which causes a significant reduction of the control effectiveness. The results show that the stability augmentation system can significantly improve the longitudinal controllability of an aircraft with a morphing wing.

This study is necessary to analyse the effect of a morphing wing on an UAV and perform a comparison with the rigid model.

The control surfaces assignment plan for trim, pitch and roll control was obtained. An active control algorism for the morphing wing was created to satisfy the required stability and control effectiveness by operating the LE and TE control surfaces according to flight conditions. The aeroelastic effect of control derivatives on the morphing aircraft was considered.

The research work presented in this paper deals with the design of a reconfigurable flight control system (RFCS) based on the control distribution concept (CDC). The work presented here is concerned with the reconfiguration in response to single and multiple control surface failures. A model of a generic fighter aircraft possessing a relatively large number of control surfaces was selected as a test model for the reconfiguration algorithm. The failure in the form of a control surface stuck at a non‐neutral position is one of the most performances degrading and challenging to recover from by a process of reconfiguration. An algorithm was developed to reconfigure the overall FCS by way of redistributing the control effort to the remaining healthy control surfaces, while at the same time cancelling the effects of control surface deflection locked at a non‐neutral position by the generation of another compensating signal feeding to the FCS.

In more than 100 years of aviation, significant progress has been made in flight control systems. The aircrafts that have entered service for the past ten years tend towards power-by-wire flight control with electrical actuators. The purpose of this study is to analyse the effects of electrical actuation on power consumption, weight and fuel consumption on a commercial transport aircraft.

The Airbus A321-200 aircraft was chosen as a case study for analysing the effects of electrical actuation on the flight control actuation system (FCAS) architecture, and Pacelab SysArc software was used for design, modelling and analysis. As alternatives to the existing system, hybrid and all-electric models are built to a set of design guidelines with certain limitations.

Compared to the existing FCAS architecture model, 80 kg weight savings in the hybrid FCAS architecture model and 171 kg weight savings in the all-electric FCAS architecture model were observed. In terms of fuel consumption, it has been observed that there is 0.25% fuel savings in the hybrid FCAS architecture model, and 0.48% fuel savings in the all-electric FCAS architecture model compared to the existing FCAS architecture model at 3200 NM.

In line with the data obtained from this study, it is predicted that electrical actuation is more preferable in aircraft, considering its positive effects on weight and fuel consumption.

In this study, three different models were created: the existing FCAS architecture of a commercial transport aircraft, the hybrid FCAS architecture and the all-electric FCAS architecture. Hybrid and all-electric models are built according to a set of design guidelines, with certain limitations. Then, similar flight missions consisting of the same flight conditions are defined to analyse the effects of power consumption, weight, and fuel consumption comparatively.

The pilot's control member 10 of an aircraft controls two power units 25, 125 which each execute movements corresponding in amount and direction to the movement of the control member and thus, through a differential gear 31, drive a torque shaft 32 which drives the irreversible gearing 50 associated with the control surface, and means are provided to prevent reverse rotation of one power unit, should it fail, by the other. The control member 10 is connected by a chain drive 11, shaft 12 and chain drives 13, 14 to two selsyn transmitters 16, 116 adapted to control the power units 25, 125 respectively. Each power unit comprises a constantly rotating constant speed electric motor 26 driving a variable delivery pump 27 which provides pressure fluid to drive a hydraulic motor 28, and each includes a mechanical lock normally held out of engagement by fluid pres‐sure, but spring‐loaded to lock the motor should it fail. The pump 27 normally docs not deliver fluid, but the delivery is controlled by a lever operated by follow‐up gears 33 in turn controlled by the output of the motors 28 and by the selsyn receiver 29 (see also Group XXIX) to provide a correspondence control. The pumps 27 and motors 28 are of the variable stroke rotary type to provide control in two directions of rotation. The motors 28 each drive one sun‐wheel 30, 130 of a differential gear 31, the planet carrier of which drives the torque shaft 32, and also the shafts 65 associated with the follow‐up gears 33 through gearing 55, 56, 64. The torque shaft 32 drives, through bevel gearing and shafting, jacks 50 each comprising a worm 49 engaging a worm‐wheel on a nut 75 engaging a screw 76 to move the surface 24. The surface 24 is in independent sections, each having its own jack, and the worms 49 are secured to the driving shaft 46 by frangible pins which break if the corresponding sec‐tion of the surface jams, so permitting continued operation of the rest of the surface. ‘Desynn’ (Registered Trade Mark) transmitters 34, 37, associated with the control member 10 and this surface 24 respectively, control indicators 35, 36 in the pilot's cabin. A device 20 for simulating ‘feel’, such as described in Specifications 619,987 and 619,988, and a trimming device 21 both operate through a differential gear 18 and chain drive 17, while the automatic pilot is connected to the shaft 12. A further differential gear associated with the shaft 32 may provide for operation of the surface by a reversible electric motor if both power units 25 fail. In an aircraft with a pressure cabin 114, the torque shaft 32 passes through a seal 45, the jacks 50 being outside the cabin and the power units 25 inside, but not necessarily in the pilot's compartment. Specification 577,496 is referred to.

THE importance of control surface mass balancing does not need to be emphasized, and it is well known that the faster aeroplanes fly, the greater is the care and attention which the designer must give to this question.

The purpose of the paper is to present an innovation-based new actuator/surface fault detection and isolation (FDI) method, which is sensitive to the changes in the innovation mean of the Kalman filter (KF) and the KF tuning method for the case of actuator/surface failure.

The multiple system noise scale factors (MSNSFs) are used in this method as the monitoring statistics. MSNSFs are determined to make it possible to perform the actuator/surface FDI operations simultaneously.

The introduced FDI algorithm can detect and isolate the loss of effectiveness type actuator/surface faults in real time. The proposed KF tuning method works effectively against actuator/surface fault. The actuator/surface fault detection, isolation and filter tuning are achieved by just using a simple modification over the conventional KF.

The MSNSF-based actuator/surface fault detection, isolation and filter tuning algorithms are investigated together for the first time. The actuator/surface FDI operations are performed simultaneously.

The purpose of this paper is to present a method for determining the safe flight boundaries of the asymmetrically loaded airplane in the terminal flight phases. The method is applicable to both, the inherent airplane asymmetries and those asymmetries resulting from the airplane use irregularities, asymmetric stores under the wing being one of the examples. The method is aimed to be used in the airplane design and combat airplane service life support.

The analysis method is based on the comparison of demanded and structurally available flight control displacements. Control surface aerodynamic properties, structurally available flight control displacements and dynamic pressure define control surface authority as the capability of control surfaces to generate the forces and moments needed by the airplane to perform required maneuvers. Demanded flight control displacements are those related to the maneuvering requirements and to those needed to compensate lateral wind and any type of the asymmetric airplane load.

The method results are given in the form of the speed and lateral wind component and are a subset of the total set of airplane safe flight boundaries. The key objective is the improvement of flight safety of the asymmetrically loaded airplane.

The method supplements the safe flight boundaries of the symmetrically loaded airplane, the minimal landing speed being the dominant limitation. This boundary positions method analysis in the domain of linear lift coefficient variation, as the function of the angle of attack permits the addition of control surface displacements required to perform the maneuvers and compensate the asymmetrical loads.

The method combines a simple roll dynamics model, stationary equations of the airplane lateral-directional motion and several numeric analysis procedures to obtain the results. This new combination possesses synergy properties and is implemented as the computer program.

This method will become a new development trend in subgrade structure design for high speed railways.

This paper summarizes the structural types and design methods of subgrade bed for high speed railways in China, Japan, France, Germany, the United States and other countries based on the study and analysis of existing literature and combined with the research results and practices of high speed railway subgrade engineering at home and abroad.

It is found that in foreign countries, the layered reinforced structure is generally adopted for the subgrade bed of high speed railways, and the unified double-layer or multi-layer structure is adopted for the surface layer of subgrade bed, while the simple structure is adopted in China; in foreign countries, different inspection parameters are adopted to evaluate the compaction state of fillers according to their respective understanding and practice, while in China, compaction coefficient, subsoil coefficient and dynamic deformation modulus are adopted for such evaluation; in foreign countries, the subgrade top deformation control method, the subgrade bottom deformation control method, the subsurface fill strength control method are mainly adopted in subgrade bed structure design of high speed railways, while in China, dynamic deformation control of subgrade surface and dynamic strain control of subgrade bed bottom layer is adopted in the design. However, the cumulative deformation of subgrade caused by train cyclic vibration load is not considered in the existing design methods.

This paper introduces a new subgrade structure design method based on whole-process dynamics analysis that meets subgrade functional requirements and is established on the basis of the existing research at home and abroad on prediction methods for cumulative deformation of subgrade soil.

Describes preliminary structural design work on a notional uninhabited tactical aircraft (UTA), carried out at Cranfield University. UTAs are seen as an important future element of military fleets. A notional baseline requirement was derived, leading to the evolution of a design solution. The basic requirements for such a UTA are naturally highly classified but, although industry has been hesitant to comment, the baseline requirements and design solution developed herein are believed to be reasonable.