Application description

The need for the MEA arose directly from the global need to minimise carbon dioxide (CO2) released during the combustion cycle of carbon based fuels. The Kyoto agreement of 1997 was the first treaty between countries to minimise global emissions of CO2. The direct effect

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of the treaty on Europe alone was to reduce the collective emissions of CO2 to 8 % below the 1990 levels by 2012 according to the European commission for climate action[1], this was further extended in 2007 to commit to a 20 % reduction by 2020. The agreement made in Cancún during 2010 marked an important step to making a comprehensive and legally binding framework to legislate beyond 2012, the end date of the Kyoto agreement. A report edited by Krzyzanowski[2] on behalf of the World Health Organization (WHO) explained the effects on health of transport related air pollution which reinforced the need to reduce the use of carbon based combustion fuels. In the United Kingdom the government [3] have stated that in 2012 approximately 6.4% of the main greenhouse gas (CO2) emissions from the U.K was directly contributed by aircraft and by 2020 this was expected to be 10 % due to the increased demand for air travel.

The MEA, like all aircraft, would combust carbon fuels for propulsion. In previous aircraft, developed adhoc over decades, the energy from the combustion was also converted in to four other forms: pneumatic, hydraulic, mechanical and electrical energy as described by Rosero[4]. The aims of the MEA were to gain fuel efficiencies through removal of all conversions other than that of converting to electrical energy via the use of generators. Indeed, the MEA program run by the U.S. Air Force aimed to deliver a Power Optimized Aircraft (POA) which reduced non propulsion power, to improve fuel efficiency, while increasing the reliability and safety of on board systems and reducing maintenance costs, again as described by Rosero[4]. The targeted reliability improvements of MEAs as

compared to conventional aircraft technology were between 1400 to 1900 % while offering a 200 % improvement in power density as discussed by Cloyd[5].

The changes necessary to the power distribution system for the MEA meant that increased power distribution voltages would be required in order to limit conduction losses, reduce cable size and hence weight, as described by Avery[6]. The targeted distribution voltage was +/- 270 V dc (load voltage =540 V) as described by Avery. In the event of generator failure the Auxiliary Power Unit (APU) would supply power again at +/- 270 V

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dc. It is likely that some form of energy storage would be used as the APU, one example being the hybrid fuel cell and Li-ion batteries as described by Eid[7]. Such Li-ion batteries have high energy storage and can reach extremely high short circuit currents (measured at a peak of 11,697 A with the full cell capacity being delivered within a three minute duration using a 50 Ah battery) as described by Issacs[8].

The power management and distribution system development for an MEA project in the U.S. was described by Maldonado[9]. Maldonado described that the power distribution system was required to not only minimise weight and maximise supply power availability, but also to maximise reliability and therefore have high fault tolerance. The power switching element was the most essential part of the power distribution network. The essential purpose of the switch was to protect the hundreds of kilometres of insulated wire in order to prevent loss of power, loss of loads and reduce fire hazard as described by Bailey[10]. The critical nature of the aircraft application meant that safety and reliability was the key factor, as described by Bailey[10], such that:

a) No single failure shall lead to a loss of power channel b) Fail safe switch node mechanisms were required c) No cascading failures allowed

d) No common mode failures allowed e) Must be designed for 24 year service life

f) Must operate reliably in harsh environments: EMC, vibration, acoustic noise Each switching node, called an Electrical Load Management Centre (ELMC) by Maldonado, consisted of a Solid State Power Controller (SSPC) and a microprocessor controlled fault detection system which provided a programmable overload curve (I2t) and current limiting, plus a fast acting short circuit trip (10 μs), again as described by

Maldonado[9]. Thus, any switching element used in the SSPC had to withstand fault current conduction for a minimum of 10 µs.

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The SSPC shown in figure 2.1 was described by Friedman[11] and utilised solid state switching devices instead of older Contact Breaker (CB) technology. The

Figure 2.1. Two examples of solid state SSPC units as supplied by Data Device Corporation

[11, 12].

specifications or CB technology, such as the EV200 series, as manufactured by Tyco Electronics Corp[13], demonstrated the main problem concerning reliability of the CB, or electro-mechanical contactor. The CB was guaranteed to break only above 650 A, and also required replacement in the event of breaking a high magnitude fault current of >2000 A (at 270 V) as shown in figure 2.2. It was also possible for the contacts to weld shut due to arcing if the current exceeded that specified by the manufacturer, for example if Li-ion batteries were used then short circuit current could be >5 times larger than the Tyco maximum break current specification as described by Issacs[8]. However, CBs were still the preferred solution at nominal currents of 50 A to 500 A as described by Bailey[10] when speaking in 2009 due to the very low on state resistance (where on state resistance, RON = 0.2 mΩ in the

EV200 series). In addition CBs had a high power density and were proven to be robust to high temperature operation and voltage transients; they also offered good isolation as provided by the air gap within the CB. Unfortunately, CBs were slow to switch, they required additional electronics to provide protection from high current, and they offered no means to control the current magnitude, again as described by Bailey[10].

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Another important reason for requiring solid state replacement of CB technology was provided by Friedman[11] and Simon[14] in that computer control of the power distribution system enabled remote placement of the ELMC between generator and loads directly, instead of routing heavy cabling via the flight engineers fuse panel. Status of the EMLC could then be relayed back to the control computer in real time due to the speed

Figure 2.2. a) A typical electro-mechanical contactor; b) and c) associated switching cycle life data reproduced from manufacturers data.

of the solid state devices. Bailey[15] agreed that the main advantage of the solid state ELMCs was that they could be positioned in remote locations directly between the bus and the load to form an overall distribution network, the optimisation of which could be tailored to each aircraft. Indeed Bailey went on to say that the SSPC solution offered a means to the provision of enhanced diagnosis and prognosis of problems. Such SSPC enabled distribution networks at +/- 270 V dc could be modelled, as demonstrated by Izquierdo[16]. The work of Izquierdo demonstrated the ability of the software tools to optimise and verify any layout, thus highlighting any potential issues through both secondary and primary switch nodes in the event of changing a load or SSPC type for example.

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The solid state switching devices were beneficial in a computer controlled system due to the greatly reduced the switching time (toff = 12 ms in the case of the EV200 series),

which eliminated contact bounce as described by Barrado[17], thus providing a reliable high fault current turn off capability. The solid state device solution to the needs of a remote ELMC however, raised a few new problems. As described by Simon[14] the solid state devices available to an SSPC designer could never match the low on state resistance of the contact breaker (CB). Due to the requirement for the use of solid state devices with a minimum BV > 500 V (from MIL-STD-704-F[18]) then increased junction temperature within the semiconductor, relative to a lower BV device, were anticipated but, according to Simon[14] the airframe had no capability to remove heat generated from excess power dissipation easily, power dissipation therefore needed to be minimised to reduce self heating. Unlike the CB solution, according to Simon[14] and Bailey[10], the SSPC could not provide galvanic isolation therefore leakage levels from the solid state power devices was required to be very low (<1 mA), even at high temperatures.

Finally, the electrical loading supplied via an SSPC would predominately be inductive in nature as described by Weimer[19]. The switching device therefore would need to achieve low static conduction losses, a high current switching capability and demonstrate lowest dynamic losses during that switching event to prevent device failure for reasons of transient lattice heating.