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Under a licence agreement negotiated with Bristol Aircraft Ltd., Haven Engineering Co. Ltd., of 29 Corsica Street, Highbury, London, N.5, are to manufacture and market the Bristol thermionic valve mounting clip.

IN this article it is proposed to deal with the broad principles of operation of the various systems of control of aircraft electrical generators which are in common use, rather than with the detail design features and the many minor variations adopted by different manufacturers of this type of equipment.

The purpose of this paper is to implement a closed loop regulated bidirectional DC to DC converter for an application in the electric power system of more electric aircraft. To provide a consistent power supply to all of the electronic loads in an aircraft at the desired voltage level, good efficiency and desired transient and steady-state response, a smart and affordable DC to DC converter architecture in closed loop mode is being designed and implemented.

The aircraft electric power system (EPS) uses a bidirectional half-bridge DC to DC converter to facilitate the electric power flow from the primary power source – an AC generator installed on the aircraft engine’s shaft – to the load as well as from the secondary power source – a lithium ion battery – to the load. Rechargeable lithium ion batteries are used because they allow the primary power source to continue recharging them whenever the aircraft engine is running smoothly and because, in the event that the aircraft engine becomes overloaded during takeoff or turbulence, the charged secondary power source can step in and supply the load.

A novel nonsingular terminal sliding mode voltage controller based on exponential reaching law is used to keep the load voltage constant under any of the aforementioned circumstances, and its performance is contrasted with a tuned PI controller on the basis of their respective transient and steady-state responses. The former gives a faster and better transient and steady-state response as compared to the latter.

This research gives a novel control scheme for incorporating an auxiliary power source, i.e. rechargeable battery, in more electric aircraft EPS. The battery is so implemented that it can get regeneratively charged when primary power supply is capable of handling an additional load, i.e. the battery. The charging and discharging of the battery is carried out in closed loop mode to ensure constant battery terminal voltage, constant battery current and constant load voltage as per the requirement. A novel sliding mode controller is used to improve transient and steady-state response of the system.

More electric aircraft (MEA) is defined as the extensive usage of electric power in aircraft. The demand for electric power in new generation aircraft rises due to environmental and economic considerations. Hence, efficient and reliable starter/generators (SGs) are trending nowadays. The conventional main engine starting system and power generation system can be replaced with an individual SG. The constraints of the SG should be investigated to handle the aviation requirements. Even though the SG is basically an electric machine, it requires a multidisciplinary study consisting of electromagnetic, thermal and mechanical works to cope with aviation demands. This study aims to review conventional and new-generation aircraft SGs from the perspective of electric drive applications.

First of all, the importance of the MEA concept has been briefly explained. Also, the historical development and the need for higher electrical power in aircraft have been indicated quantitatively. Considering aviation requirements, the candidate electrical machines for aircraft SG have been determined by the method of scoring. Those machines are compared over 14 criteria, and the most predominant of them are specified as efficiency, power density, rotor thermal tolerance, high-speed capability and machine complexity. The features of the most suitable electrical machine are pointed out with data gathered from empirical studies. Finally, the trending technologies related to efficient SG design have been explained with numeric datasets.

The induction motor, switched reluctance motor and permanent magnet synchronous motor (PMSM) are selected as the candidate machines for SGs. It has been seen that the PMSM is the most preferable machine type due to its efficient operation in a wide range of constant power and speed. It is computationally proven that the using amorphous magnetic alloys in SG cores increases the machine efficiency more. Also, the benefits of high voltage direct current (HVDC) use in aircraft have been explained by a comparison of different aircraft power generation standards. It is concluded that the HVDC use in aircraft decreases total cable weight and increases aircraft operation efficiency. The thermal and mechanical tolerance of the SG is also vital. It has been stated that the liquid cooling techniques are suitable for SGs.

The demand for electrical power in new generation aircraft is increasing. The SG can be used effectively and efficiently instead of conventional systems. To define requirements, constraints and suggestions, this study investigates the SGs from the perspective of electric drive applications.

WITH THE RAPID GROWTH of world airline investment in powered ground equipment, much closer attention is being given to the development of fixed rather than mobile ramp services, for electrical, air starting, lavatory and water servicing of aircraft. Mobile Ground Power Units account for a large proportion of the capital investment, often at the highest unit cost.

The purpose of this paper is to focus on the machine design and control strategy of the permanent magnet synchronous generator (PMSG) system, especially utilized in variable speed applications, in order to stabilize the output voltage on the dc link over a wide speed range.

Different ac/dc power converter topologies are comparatively studied, each with an accordingly designed PMSG, so as to investigate the influence of the armature winding inductance as well as the relationship between the PMSG and power converter topologies.

Pulse width modulation (PWM) rectifier is preferable for the said application due to its good performance and controllability. Moreover, by employing the PWM rectifier, relatively large inductance of the PMSG is considered for both short-circuit current reduction and field regulation.

Field-regulating control is realized with a space vector PWM (SVPWM) rectifier, which can weaken the PMSG magnetic field during high-speed operation, while even properly enhance the field at low speed, ensuring a small change of the PMSG output voltage and a stable dc voltage.

AS an aircraft flies through the atmosphere it is heated kinetically owing to its forward velocity. For subsonic speeds the effect is hardly measurable, at M= 1 ·5 it is definitely measurable and as there is a rapid change with speed it soon becomes necessary to include kinetic heating in the design conditions for an aircraft structure. All other conditions that are present at lower speeds have still to be retained and it becomes a matter of adding heating conditions to an already large number of conditions. The same approach must be used of preparing a design envelope on which appropriate factors have to be applied. In doing this it should be appreciated that there are two distinct effects of heating, one is the steady temperature condition associated with sustained steady flight conditions, the other is the rapid change in temperature and associated structural stresses and distortions when the aircraft changes speed or height. Considering first the steady temperature condition, it is evident that this can only arise in practice after a fairly long time at the particular flight condition to which it applies and that intermittent departures from it will not have a significant effect. The aircraft speed that has to be selected must of course be one that might reasonably be expected to be sustained occasionally for moderate periods, although perhaps not quite long enough to reach equilibrium. There is a comparable case in the normal strength requirements for gusts. The design gust has to be associated with an appropriate aircraft forward speed namely ‘Design Cruising Speed’. It is suggested that exactly the same speed be used to determine the steady temperature conditions with no further safety factor, and that all static and fatigue strength conditions be satisfied with full safety factors at this temperature condition.

THE electrical supply system is designed to provide both 200 volt a.c. and 28 volt d.c. services for operation of the aircraft. Although electrical power is required for the operation of the auto‐stabiliser system, this is not essential for flight and the flying controls are non‐electric. The aircraft does not therefore, depend on the integrity of the electrical system for flight. Loads, such as undercarriage lowering selection etc., required to operate if a total generation failure has occurred, are all d.c. and are fed by the main aircraft batteries. As a further precaution basic flight and communication facilities can also be powered in an emergency from a standby battery which is not charged in flight and is therefore independent of any of the normal aircraft services.

This article is aims to inform aircraft propulsion system designers of the implications which fundamental power distribution design assumptions have on the effectiveness and viability of turboelectric distributed propulsion (TeDP) systems. Improvements and challenges associated with selecting alternating or direct current for normal- and superconducting distribution systems are presented. Additionally, for superconducting systems, the benefits of bi-polar DC distribution are discussed, as well as the implications of operating voltage on the mass and efficiency of TeDP grid components.

The approach to this paper selects several high-level fundamental configuration decisions, which must be made, and it qualitatively discusses potential implications of these decisions.

Near term TeDP architectures which employ conventionally conducting systems may benefit from alternating current (AC) distribution concepts to eliminate the mass and losses associated with power conversion. Farther term TeDP concepts which employ superconducting technologies may benefit from direct current (DC) distribution to reduce the cryocooling requirements stemming from AC conduction losses. Selecting the operating voltage for superconducting concepts requires a divergence from the present day criteria employed with terrestrial superconducting transmission systems.

The criteria presented in the paper will assist in the early conceptual architecting of TeDP systems.

The governing principles behind the configuration of multi-MW airborne electrical microgrid systems are presently immature. This paper represents a unique look and the motivating principles behind fundamental electrical configuration decisions in the context of TeDP.