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The safety critical electric machines and drives in the more electric aircraft-A survey


The Safety Critical Electric Machines and Drives in the More Electric Aircraft: a Survey
A. Boglietti, Senior member, IEEE, A. Cavagnino, Member, IEEE, A. Tenconi, Member, IEEE, and S. Vaschetto Politecnico di Torino, Dipartimento di Ingegneria Elettrica, Corso Duca degli Abruzzi 24, 10129 Torino, ITALY aldo.boglietti@polito.it, andrea.cavagnino@polito.it, alberto.tenconi@polito.it, silvio.vaschetto@polito.it
Abstract - In order to improve aircraft efficiency, reliability and maintainability, the aerospace world has found in the progressive electrification of on-board services the way to reduce or to remove the presence of the hydraulic, mechanical and the bleed air/pneumatic systems. This concept is called More Electric Aircraft (MEA), which involves the introduction of the Electromechanical actuators (EMAs) and the electrohydraulics actuators (EHAs) for the actuation of the flight surfaces of wide-body aircraft, moving from the “fly-by wire” to the “power-by wire” concept. The resulting step change in aircraft electrical loading will have far reaching implications for the electrical generation system. Considerable effort is being directed towards realizing the so-called More Electric Engine (MEE), which foresee an integration of the electrical generator directly inside the main gas turbine engine. Also the entire electrical distribution system is subject to radical revisiting, with a trend which is leaving the constant frequency AC energy distribution in favor of variable frequency or DC solutions. Hence, it is evident the MEA trend increments the technical challenges and research topics for the electrical engineering in the aeronautic applications. The aim of the paper is the presentation, from the electrical engineering point of view, of some of the most challenging application of electric machines and drives in the incoming new aircraft generation. There are too many on board electric components and systems to be analyzed: the authors centre the attention on the most safety critical drives: flight surface actuators, fuel pumps and generators.

I.

INTRODUCTION

Keywords: more electric aircraft, more electric engine, safety critical drives, electric motors, electric generators, actuators. LIST OF THE MAIN USED ACRONYMS AEA APU BLDC CVG EBHA EHA GCU HP IAP IDG JAA LP MEA MEE MESA MESEMA MOET POA RAT All Electric Aircraft Auxiliary Power Unit Brushless Direct Current Motor Constant Velocity Gearbox Electrical Backup Hydraulic Actuators Electric-Hydraulic Actuators Generator Control Unit High Pressure Integrated Actuator Package Integrated Drive Generator Joint Aviation Authority Low Pressure More Electric Aircraft More Electric Engine Magnetostrictive Equipment and Systems for more electric Aircraft (project) Magnetoelastic Energy System for Even More Electric Aircraft (project) More Open Electrical Technologies (project) Power Optimised Aircraft (project) Ram Air Turbine

Technological advances in the aircraft industry have improved aircraft efficiency and reduced the costs of air transport by such a degree that worldwide air passenger traffic has grown at an average yearly rate of 9% since 1960; today it is postulated that passenger air traffic will grow at a rate of between 5 and 7% into the foreseeable future, and even faster will be the growth of cargo traffic. Today air transport produces 2% of man-made CO2 emissions, this is expected to increase to 3% by 2050. In this contest, there are many environmental as well as commercial pressures on aircraft manufacturers to improve the performance of future aircraft in terms of safety, air pollution, noise and climate change. To achieve these goals it is necessary revisiting the whole aircraft architecture system, with the introduction of new technologies for performing key functions on aircraft. Today the conventional civil aircraft are characterized by four different secondary power distribution systems: mechanical, hydraulic, pneumatic and electrical. This implies a complex power distribution nets aboard, and the necessity of an appropriate redundancy of each of them. In order to reduce this complexity, with the aim to improve efficiency and reliability, the aircraft manufacturer trend is towards the More Electric Aircraft (MEA) concept that is the wider adoption of electrical systems in preference to the others. The resulting step change in aircraft electrical loading will have far reaching implications for the electrical generation systems, realising the so-called More Electric Engine (MEE), in which the electrical machines are integrated inside the main gas turbine to generate electrical power, start the engine and guarantee safety generation in case of a critical on-flight failure. In the past years many projects and initiatives have been developed to explore the MEA/MEE concepts both for military and civil applications, towards the All Electrical Aircraft (AEA). In 2000 the MESA (Magnetostrictive Equipment and Systems for more electric Aircraft) project was launched, aimed to reducing power take up and weight of on-board aircraft systems through the development of magnetostrictive motors and actuators. In 2002 POA project (Power Optimised Aircraft) was aimed to the validation at aircraft level and both qualitatively and quantitatively, the ability of alternative

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Fig. 1. Wing control surfaces of a fixed-wing aircraft: 1. wingtip, 2. low speed aileron, 3. high speed aileron, 4. flap track fairing, 5. Krüger flaps, 6. slats, 7. three slotted inner flaps, 8. three slotted outer flaps, 9. spoilers, 10. spoilers air-brakes (source: Wikimedia Commons, [3]).

equipment systems to reduce weight, fuel consumption and maintenance costs. In 2004 the MESEMA (Magnetoelastic Energy Systems for Even More Electric Aircraft) project was devoted to the development, production and test of “innovative transducer systems based on active materials” aimed for high-torque actuation, vibration and noise reduction, electrical energy generation and structural health monitoring. This project has been evolved in two European research programs named “MADAViC” and “MESA” based on the six years scientific and technological objectives reached by the MESEMA. In 2006 the MOET (More Open Electrical Technologies) project, was aimed to establish the new industrial standard for commercial aircraft design, in conjunction with the reducing on the aircraft emission and improving the operational capacity, evolving from the “flyby-wire” to the “power-by-wire” concept. Today the MEA/MEE topics have a relevant role in the research projects managed by the CleanSky, equally shared by the European Commission and industry, over the period 2008-2013. All these projects have contributed to the development of many electric equipments that are now installed in the Airbus A380 and which will be employed in the Boeing B787, which are the today maximum expression of the MEA concept. The fundamental aspects that characterize the safetycritical systems of the aeronautical components are the stringent reliability requirements, such those established by the JAA (Joint Aviation Authority) [1]. Focusing on the electrical generation and distribution system, the power quality generated on-board must also fulfil the MIL-STD704F standards [2]. Regarding the generator integration inside the main gas turbine engine, besides the standards, not of secondary importance are the minimization of the weight and volume, as well as the harsh operating environment (especially high temperatures). All these aspects introduce a new level of complexity in the design of the electrical components dedicated at those aerospace applications. In conclusion, it is evident the paramount importance of reliability for some of the electric drives in MEA/MEE applications; in particular the paper deal with three main drives that are safety-critical:

Fig. 2. Tail of a Lufthansa Airbus A319 (source: Wikimedia Commons. Commons is a freely licensed media file repository, [3]).

? the electro-mechanical actuators for primary flight surfaces control (section II); ? the electric fuel pump (section III); ? the starter-generator embedded within the engine (section IV). The MEA concept involves also a redefinition of the onboard power distribution systems; some consideration about that are reported in section V. The more electric aircraft approach is widely discussed in the technical literature, especially in the aerospace field; the aim of the paper is the presentation of some of the most interesting research issues concerning the electric machines and drives for safety critical applications, analysed from the point of view of the electrical engineering specialist. II. FLIGHT SURFACES CONTROL

In the wings and in the tail of a wide-body aircraft there are several surfaces that the pilots can move/adjust in order to stabilize the airplane trajectory and to control the lift on the wings. Examples of these surfaces are reported in Fig.1 and Fig.2. These adjustable flight surfaces can be subdivided in two groups with respect to their main functionality: the “primary” and “secondary” flight controls. The primary flight controls (ailerons, elevator and rudder) are used to control the roll, pitch and jaw, even if they can perform secondary effects too [3]. The secondary flight controls, also called as high lift system, serves to change the wing lift. The number and type of actuators is very different, with respect to the considered aircraft. In addition, the load requirements are very different too: starting from few kilowatts for the edge slats, up to 50-60 kW for the horizontal stabilizers and the rudder [4]. Also the dynamic load profile can be quite different: there are few surface movements with very large extension and short duration (typically during the landing and take-off) or several “small” surface adjustments during the flight [5]. In addition, anomalous performances are generally requested to the actuators in critical flight conditions. Just for example, if all the engines on the same wing fail, the rudder actuator has to be able to

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Fig. 4. Example of large Electric-Hydraulic Actuator (EHA). Source: [9]. Fig. 3. High-power Electromechanical Actuator (EMA). Source: [9].

keep the rudder in a fixed position, with high yaw angle, during the flight. In this situation, very high torque is requested at the electric motor [16]. It is important to remark that the actuators have to work in very harsh ambient conditions: temperature between -60 °C and +70 °C, air pressure between almost 0 and 1 bar. Due to the low thermal conductivity of the airframe (composite materials, sheet materials, etc.), the thermal exchanges between the actuators and the surrounding environment has to be carefully evaluated [4]. In conventional wide-body aircraft, the actuation system of the flight surfaces is realized by a centralized hydraulic system, constituted by a hydraulic pump and hydraulic motor drives positioned in the fuselage plus several fluid pipelines and hydraulic actuators positioned in the wings and tail surfaces. The control of the hydraulic actuators is realized with the well-established “fly-by wire” technology, where no mechanical links between the control surfaces and the cockpit handles are present [6], [7]. Moving towards an all-electric aircraft scenario, the idea to control each surface with an own directly coupled electromechanical actuator (EMA) is a must. This concept is defined as “power-by-wire” [7]-[9]. An example of an EMA for large flight surface is shown in Fig.3. Due to safety and reliability reasons, mainly concerning the jamming vulnerability (gearbox or ballscrew for rotary-tolinear movements), the air framers had still now some concerns to use EMAs for primary flight control surface preferring the most reliable electric-hydraulic actuators (EHA). In the EHAs there are still a hydraulic circuit, but it is just confined in each actuator to transmit power from the electric motor to the surface (Fig.4) [10]. The main advantage of an EHA is that the actuator can be controlled as a conventional hydraulic one, obtaining the traditional activestand by or active-active device operations [9]. In [9], the Integrated Actuator Package (IAPTM) is presented. This device is an EHA that, thanks to an advanced dual-channel hydraulic circuit, allows to use an unidirectional constantspeed electric motor.

When some, but not all, of the traditional hydraulic circuits are removed and substituted by EMAs and/or EHAs, it is common to speak of “more electric aircraft” (MEA). With respect to the flight controls, the first application of EHAs to primary flight surfaces was in the delta-wing Vulcan bomber in the 1950s [11]. Its redundant design, achieved using the EHAs, allowed to get an impressive safety record. More recent examples of commercial MEAs are the Boeing 787 and the Airbus A380. In the Boeing 787, a mid-sized wide-body aircraft, spoilers and horizontal stabilizer flight controls are operated by electric motors in order to guarantee the control functionality also in the case of a total hydraulics failure. The super-jumbo A380 represents the state-of-art with respect to the flight control systems. As reported in detail in [11], in the A380 aircraft many EHAs or EBHAs have been introduced in several control surfaces, allowing to get redundant power sources for the surface actuations. The EBHAs are actuators that provide backup electrical power at the surface through a local electric motor and an associated hydraulic pump. EBHAs are hydraulically powered in the normal mode and electrically powered in backup mode.

Fig. 5. Scenario of the EMA introduction in aircraft flight control systems (Power source in the vertical axis on the left: M=Mechanical, H=Hydraulic, E=Electrical; Actuator type on the right). Source: [7].

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On the basis of the previous considerations, the possibility of an electric actuation of the flight surface is beyond dispute for its potential advantages with respect to conventional hydraulic solution [9], in particular for the expected benefits in terms of overall weight reduction, better reliability and safety, reduced costs (maintenance, operational and fuel consumption); Fig.5 shows the future vision concerning the introduction of the EMAs in aircraft flight control systems [7]. As shown in Fig.5 the use of EMAs, deleting all hydraulic pumps and pipelines, is the next step that has to be done to more and more approach the “all-electric-aircraft” idea. Anyway, the maturity of this new concept has to be still proven by means of researches and applicative solutions, in particular from the safety and reliability point of view. A. Electric motors and drives for EMA or EHA applications – Remarks and literature review Both EHAs and EMAs use an electric motor and a power converter plus a control system [12]. With respect to the electric motor, a literature review reveals that several types of motor can be used, but it is shown that the Brushless DC (BLDC) and the Switched Reluctances (SR) motors are the more promising ones due to their lightweight, reliability characteristics [4], [13]-[17]. Taking into account the reliability level for flight certifications, the electric drives have to be designed including fault-tolerant capabilities. As known, the faulttolerant behavior can be done using a redundancy approach or making the device itself fault tolerant. The first approach is often used in the power converter (i.e. redundant inverter legs, separated inverter for each motor phase, control system duplication, enhanced fault diagnostic functions for the power electronic switches [9], [15], [16]), while the second one is typically adopted for the electric motor. It is commonly reported that a fault-tolerant electric motor for EMAs applications has to be guarantee: ? ? ? ? ? High torque/weight ratio High torque/ampere ratio High efficiency in the full speed range Electrical, thermal, magnetic insulation between the phases and mechanical

Also the power converter topology is discussed and analyzed in literature. The proposed solutions regard the conventional Voltage Source Inverters (VSIs) and matrix converters. The converter topology influence several aspects, such as the requested DC-link capacitor in the VSIs (with room and weight problems [19] and power quality management [5]) and the power quality issues for the matrix converters [4], [20]. It is important to remark that the possibility to have a intelligent control systems in each EMAs or EHAs allows to define new fault detection strategies, including incipient faults (thanks to testability features included in the electric drives) without loosing the actuator functionality. As a consequence, a positive feedback is expected in terms of safety, reliability and maintenance costs [21]. III. ELECTRIC FUEL PUMP

Highest value of the phase inductance (in order to limit the short circuit currents)

? Safe operation in faulty conditions (one phase loss) These characteristics can be obtained both with the SR machines and the BLDC ones. Examples of surface-mounted permanent magnet BLDC motors with winding wound around a single tooth ables to verify the previous reliability requirements were proposed in [14] and [18]. The electric drive has to be designed in accordance to the selected electric power generation strategy: constant or variable frequency electric supply [11].

The fuel pumps can be subdivided in two categories: the low pressure boost / transfer pump types and the high pressure FCU (Fuel Control Unit) fuel pump. The first pump category is normally electrically operated, while, in traditional systems, the high pressure fuel pump is directly driven through the mechanical gearbox and the fuel flux is controlled by means of the fuel valve. As a consequence, the focus is on the high pressure fuel pump because it is another aircraft apparatus that could be electrically driven, introducing the concept of “smart electric fuel pump”. The main advantage of the electric solution for the fuel pump is in the possibility to drive the pump at variable speed. In this way the pump can deliver a variable fuel flux to the combustion chamber, in accordance to the engine control requirements, eliminating the fuel valve in the fuel metering system [8]. In [18] an analysis of the required fault-tolerant behaviors for electric drives used in fuel control system are discussed and multi-phase PM motor prototypes are considered. The same authors describes a four phase PM machine specifically designed for an engine fuel pump in [22], while they analyzed the faulty drive operations in [23]. The electric motor design constraints, imposed by reliability and fault-tolerant requirements, are the same ones reported in section I.A. Some engine actuators of the fuel control system can be arranged as “more electric” too. For example, the feasibility of “smart electric fuel valve” able to control the fuel flux was reported in [8] and [11]. The application of this new technology leads to several advantages, such as weight saving, lower maintenance costs and improved in-service reliability. IV. ON-BOARD POWER GENERATION

On conventional civil aircraft, the electrical power is usually generated by wound field synchronous generator with a PM exciter stage [24] [25]. A Generator Control Unit (GCU) performs a field control in order to regulate the terminal voltage.

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The generator is mechanically driven by the main engine shaft by means of a Constant Velocity Gearbox (CVG), allowing to maintain constant the frequency at 400 Hz. If the CVG is integrated inside the generator, it is called Integrated Drive Generator (IDG) [26]. In Fig.6 it is depicted an example of gearbox.

use of the actuator regeneration will require a whole distribution system redefinition able to accept and manage the recovered energy. A. More Electric Engine (MEE) With the MEA concept, the electric power requirement on aircraft aboard is continuously increasing, with estimation of more than 500 kW per engine in the future [25],[30]-[32]. To reduce the general system complexity, failure probabilities and with the aim to increase the general system efficiency there are several studies devoted to the integration of the electrical generators directly inside the main gas turbine engines. This concept is called More Electric Engine (MEE). In this way the CVG (or IDG) systems could be, partially or at all, eliminated. As a consequence the generated fundamental frequency changes over a wide range in function of the engine speed variation, depending on the engine throttle [25], [30], [33].

Fig. 6. Example of gearbox. Source: [26].

In addition at the previous described energy generation systems, Auxiliary Power Units (APUs) are presents on the aircraft. They are small fuel burner jet engines connected to dedicated electrical generators, aimed to supply vital loads in case of main engines or generators failure. They are also employed to provide electric power in the pre-flight conditions, when the main aircraft engines are still turned-off. As back-up energy generation system, in addition at the APUs, there are also the Ram Air Turbines (RATs) which are propellers spanned by the high speed of the air flows near the airframe body (Fig.7). They are extracted by the airplane body only in emergency conditions.

Fig. 8. Aircraft turbofan engine (source: Wikimedia Commons, [32]).

Fig. 7. Example of Ram Air Turbine. Source: [27].

An interesting additional system of onboard power generation concerns the regeneration possibility by the electric actuators. This is possible when the energy from the loaded flight surfaces can be sent backwards to DC link using bi-directional power converter. Obviously, the energy amount depends on the load profile and actuator duty cycle. Today this energy is dissipated in resistor banks, with unavoidable weight and heat dissipation problems [27]. In the future, the

With reference to the turbofan engine structure (Fig.8) it is possible to integrate the generator inside the main engine in some different positions. In particular, the generator can be driven both by the Low Pressure (LP) and the High Pressure (HP) shaft. These two possibilities involve different advantages and disadvantages, mainly concerning dimensions, speeds and environmental working conditions. When the LP shaft integration is selected, the generator is characterized by a lower rotational speed, but higher radial dimensions to achieve the same rated power with respect to the HP shaft integration solution. However, the LP shaft integration guarantees better environmental conditions, especially regarding the ambient temperatures. It is important to highlight that LP shaft connects the low pressure compressor, turbine and the inlet fan. As a consequence, with an appropriate design of the electrical machine driven by the LP shaft, the windmill effect can be exploited [35]. In this way is possible to generate electricity in case of a catastrophic engine failure, removing the actual ram-air systems and its high maintenance costs [31]. However, since the fan rotational speed during the windmill is low, the required working speed range of the generator is very challenging (around 12-14:1, [25]).

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In the HP shaft integration, the electric machine is characterized by a lower weight and it takes up a lower room, due to the higher rotational speed of the shaft. In addition, this solution allows to use the electric machine as engine starter, avoiding in this way the pneumatic auxiliary dedicated system. Anyway, due to the high inertia of the HP gas turbine, a large torque motor capability is required at zero speed too [25], [35]. The main drawback of the HP shaft integration is the harsh environmental working conditions, mainly due to the high ambient temperatures. B. Electric generators for MEE applications – Remarks and literature review The most common electrical machines proposed in literature for MEE applications, are the Switched Reluctance (SR) machine and the Permanent Magnet (PM) ones. Some valuable examples and comparisons of these machine types are reported in [30], [31], [35], [36]-[39]. The switched reluctance machines are characterized by an intrinsic high fault tolerance, high ruggedness and construction simplicity. Another important aspect that make this kind of machine interesting for this application, is the possibility to use a single-slot coil pitch winding structure. In this way the stator coil overhangs length are very short and the bobbins results electrically insulated [36]. The main disadvantages of the SR machines are their usually lower power and torque density respect to the PM machines, high ventilation losses, small airgap, and the necessity of a more complicated power converter. Regarding the PM machines, they can be designed in several ways: with surface mounted magnets (often using Halback array PM arrangement), flux concentrating geometries, radial or axial flux topologies, inside-out radial structure, etc. The PM machines are characterized by high volumetric and gravimetric power density, very small losses in the rotor, (with the consequent cooling facilities), and high pole number (usually realized with fractional-slot windings in order to reduce the endwinding length and to get the phases decoupling). The main disadvantage of this kind of electrical machines is its unavoidable intolerance to high temperatures, due to the PMs presence. An important aspect of the PM machines is its intrinsic permanent flux, which can not be shut down in case of fault. In the literature other machine types are considered and analyzed for this application, such as induction motors [36], and special hybrid machines [30] [31] [40]. Induction motors are relatively rugged, but they are characterized by lower power density with respect to SR and PM machines. The hybrid structures presented in the references are realized with a two-part rotor, composed by a surface mounted PM and a variable reluctance section. In these prototypes an high direct-axis inductance and an improved machine’s torque capability at low speed (due to the additional reluctance torque) are obtained. A high value of

direct-axis inductance allows to obtain a constant-power speed range regulation using field-weakened strategies [31]. This characteristic is very interesting because the speed of the LP shaft is not constant (with 3:1 variations range) [35]. Independently of the selected machine type, the literature review shows an interest in ring shape motors (low axial core length/diameter ratio) in order to accommodate the geometry constraints inside the jet engine. Some applicative solutions of MEE come from the military aviation. For example, in [41]-[43] is presented a fighter aircraft with two main engines plus a third auxiliary one, for a total installed power generation system of 750 kW. In this case, a 250 kW (with a peak of 330 kW for 5 seconds), 270 VDC, switched reluctance starter/generator integrated in the gas turbine engine is presented. V. POWER DISTRIBUTION SYSTEMS

In order to achieve a fully optimized all-electrical aircraft, the whole electrical distribution system architecture should be redefined too. Nowadays, in many cases, there are two main distribution power buses on-board of conventional aircrafts [25] [30] [44]: (1) a high power, three phase, 115 V, 400 Hz devoted to large loads supply; (2) a low power, 28 VDC, for avionics and battery-driven services. Since the IDG removal is a must in the All Electric Aircraft (AEA) concept, the industry trend seems to be towards an AC variable frequency generation system, with a DC high voltage distribution bus [25]. The variable frequency strategy (called sometimes “frequency wild” [45]) does not require the IDG and as a consequence, a higher system power density is possible. Taking into account that the generators can be driven by shafts with very different rotational speeds, it is reasonable to convert all the generated power into an unique high-voltage DC distribution system output around the airframe. A high-voltage distribution system allows to reduce the cables weight because the current is lower. Moreover, the cables sizes are further reduced because in the DC system there is not the reactive power flow such as in the AC one, and there is not the skin effect due to the high current frequency [44]. In addition at the high voltage DC systems, it will remain the low voltage 28 VDC systems to supply the avionic equipments. The voltage step-down between the high voltage DC distribution system and the loads can be done in a centralized way for each load centres, as shown for example in Fig.9 and reported in detail in [44]. In the more electric wide-body aircraft, there is a transitory solution characterized by a hybrid AC and DC on-board distribution systems. For example, in the Boeing 787, which has a total power requirement of 1 MW, there is a 230 VAC at variable frequency between 360 Hz and 720 Hz, a 115

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ACKNOWLEDGMENT This work has been developed in the frame of GREAT 2020 project co-funded by Regione Piemonte. REFERENCES
[1] C. Raksch, V. Maanen, D. Rehage, F. Thielecke, U.B. Carl “Performance Degradation Analysis of Fault-Tolerant Aircraft Systems”, First CEAS European Air and Space Conference, Berlin, 10-13 September, 2007, CEAS-2007-425. MIL-STD-704F “Flight control surfaces”, Wikipedia the free encyclopedia, available http://en.wikipedia.org/wiki/Flight_control_surfaces. P.A. Robson, K.J. Bradley, P. Wheeler, J. Clare, L. de Lillo, C. Gerada, S.J. Pickering, D. Lampard, C.K. Goh, G: Towers, C. Whitley, “The impact of matrix converter technology on motor design for an integrated flight control surface actuation system”, conf. Rec. IEEE-IEMDC'03, Vol. 2, June 2003, pp.1321-1327. D.R. Trainer, C.R. Whitley, “Electric actuation-power quality management of aerospace flight control systems”, Conf. Rec. International Conference on Power Electronics Machines and Drives, June 2002, pp.229-234. P. Traverse? I. Lacaze, J. Souyris “Airbus Fly-by-Wire: A Total Approach to Dependability”, IFIP International Federation for Information Processing, Book, Springer, Boston,Vol. 156/2004, pp 191-212. T.Robin, “From “fly-by-Wire” to “Power-by-Wire” More Electric for Green Plane”, presentation at Pollutec 2008, lyon, France, 2008, available on http://www.pdfgeni.com/ref/fly-by-wire-pdf.html. TRW Aeronautical Systems, “Advances in more-electric aircraft technologies”, Aircraft Engineering and Aerospace Technology, Vol. 73, No. 3, June 2001. S.L. Botten, C.R. Whitley, A.D. King, “Flight Control Actuation Technology for Next-Generation All-Electric Aircraft”, Technology Review Journal, Millennium Issue, Fall/Winter 2000, pp.55-67. P.W. Wheeler, J.C. Clare, M. Apap, L. Ampringham, L. De Lilo, K.J. Bradley, C. Whitley, G. Towers, “An electro-hydrostatic aircraft actuators using a matrix converter permanent magnet motor drive”, IEEE-PEMD 2004 International Conference on Power Electronics, Machines and Drives, March 2004, Vol. 2, pp. 464-468. C. Adams, “A380: ‘More Electric’ Aircraft”, Avionics Magazine, October 2001, available on http://www.aviationtoday.com/av/ categories/commercial/12874.html. J. A. Weimer, “The Role of Electric Machines and Drives in the More Electric Aircraft”, Conf. Rec. IEEE-IEMDC’03, Madison, Vol. 1, June 2003, pp. 11-15. A.G. Jack, B.C. Mecrow, A. Haylock, “A Comparative Study of Permanent Magnet and Switched Reluctance Motors for HighPerformance Fault-Tolerant Applications”, IEEE Transactions on Industry Applications, Vol.32, No.44, July/August 1996, pp. 889-895. C. Gerada, K.J. Bradley, “Integrated PM Machiene Design for an Aircraft EMA”, IEEE Transactions on Industrial Electronics, Vol.55, No. 9, September 2008, pp.3300-3306. C. Cossar, L. Kelly, T.J.E. Miller, C. Whitley, C. Maxwell, D. Moorhouse, “The design of a switched reluctance drive for aircraft flight control surface actuation”, IEE Colloquium on Electrical Machines and Systems for the More Electric Aircraft, Nov. 1999, pp. 2/1-2/8. A. Garcia, J. Cusido, J.A. Rosero, J.A. Ortega, L Romeral, “Reliable electro-mechanical actuators in aircraft”, IEEE Aerospace and Electronic Systems Magazine, Vol. 23, No. 8, August 2008, pp. 19-25. P.M. Churn, C.J. Maxwell, N. Schofield, D. Howe, D.J. Powell, “Electro-hydraulic actuation of primary flight control surfaces”, IEE Colloquium on All Electric Aircraft, June 1998, pp. 3/1-3/5. B.C. Mccrow, A.G. Jack, D.J. Atkinson, J.A. Haylock, “Fault tolerant drives for safety critical applications”, IEE Colloquium on New Topologies for Permanent Magnet Machines,June 1997, pp. 5/1-5/7. M.Khatre, A.G.Jack, “Simulation of PMSM VSI Drive for Determination of the Size Limits of the DC-Link Capacitor of Aircraft Control Surface Actuator Drives”, Conf. Rec. PEDES '06, Dec. 2006, pp.1-6.

[2] [3] [4] Fig. 9. Example of load center and distribution lines for a MEA. Source: [42]. [5]

VAC at 400 Hz for components that need the traditional constant frequency supply (obtained from the previous by means of an electronic power converter), a 28 VDC bus for the avionics equipments, and a ±270 VDC (540 VDC) subbus [25] [44] [46] [47]. VI. CONCLUSIONS

[6]

[7] [8] [9] [10]

In order to improve the performance of the future aircraft in terms of safety, air pollution, noise and efficiency, a whole architecture revisiting is necessary. To achieve these goals, the military and civil aerospace world has found in the progressive electrification of on-board services the way to reduce or to remove the presence of the hydraulic, mechanical and the bleed air/pneumatic systems. Hence, the wider adoption of electrical solutions on aircrafts presents some new challenges in designing the electrical machines and drives, in particular for those applications that are safety critical. The necessities of lightness and efficiency, together with the harsh environment operating conditions, is promoting the use of advanced materials. The safety issues is promoting the use of more reliable machine structures, such as multiphase and/or concentrated windings, as well as more reliable power electronics converter topologies; at the same time prognostic and diagnostic techniques are becoming extremely important together with the control techniques for the fault tolerant management of the drives. The exigencies of maintenance in reduced accessibility area, as well as the integration requirement, is driving toward not conventional machine shapes and geometries. The high increase of power demand is pushing toward new and more articulated energy distribution network with different type of energy systems, such as wild frequency and/or “high voltage” DC. Finally, the more/all electric aircraft is one of the most powerful sources of new development in the field of the high performance electric machine and drives.

[11] [12] [13]

[14] [15]

[16] [17] [18] [19]

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