More-electric aircraft

Citation

(2005), "More-electric aircraft", Aircraft Engineering and Aerospace Technology, Vol. 77 No. 1. https://doi.org/10.1108/aeat.2005.12777aac.001

More-electric aircraft

More-electric aircraft

Keywords: Conferences, Aircraft

Aircraft have become more and more electric for many years and a recent conference at the Royal Aeronautical Society investigated the development processes by which this is coming about and looked forward to further progress in this fast moving field. Steps in the advance to date include voltage change in the 1950s, fly-by-wire (FBW) in the A320 in the 1980s, and more recently, power-by-wire electrically powered flight control. Current research programmes have great potential but problems remain to be solved, particularly with large aircraft.

The two-day conference was organized in four sessions, the first concerned with present/anticipated platforms. The initial paper was given by Lester Faleiro of Liebherr Aerospace on Trends Towards a More Electric Aircraft. Conventional power distribution includes pneumatic power bled from the engine compressor(s), a mechanical accessories gearbox, the central hydraulic pump, and the main generator providing electrical power to the avionics, etc. The road to the more- electric aircraft (MEA) is a long one and two programmes starting in the late 1990s began to look at it from an aircraft level perspective. One, the Totally Integrated More Electric Systems (TIMES) is supported by the DTI and uses previously developed systems and integrates them in an electrical network to determine the viability of using such a network in a future MEA. The other is the US Air Force Research Laboratory MEA programme which is a dual use venture but with an initial concentration on providing a more- electrical weapons capability for fighter aircraft. Another study, Power Optimised Aircraft (POA) began in 2002 and will run for 4 years.

Progress at systems level was considerable with both advantages and disadvantages becoming apparent, with more electrical systems tending to be heavier but more energy efficient and more reliable. At aircraft level the most vital achievement in the first phase of the programme has been the creation of a set of aircraft requirements. Notable results to date include the features of decreasing engine autonomy, increasing availability, the importance of snowball effects, the effects of load distribution, and the need to consolidate power electronics and drives. Overall, a civil transport MEA in the vein of POA is feasible and realisable within a surprisingly short time.

The More Electric Engine (MEE) concept was explored by Richard Newman of Rolls- Royce, whose paper described the vision of such an engine and in particular, identified the areas in which change of technology can be introduced and benefits gained from its implementation. The aircraft power Power system architecture system is heading for major changes. The Airbus A380 has an electrically powered actuator system and the Boeing 7E7 incorporates more electric technologies. Conventional aircraft engines also provide power to the aircraft systems, which includes heavy sub-systems. As part of the POA programme Rolls- Royce is investigating the MEE, which seeks to rationalise the secondary power system by replicating its functionality through electro-mechanical means.

The key design areas are high pressure starter generator (HPSG), fan shaft driven generator (FSDG) and the low pressure turbine active magnetic bearing (AMG). The HPSG is being designed by Thales and the FSDG is being designed by Goodrich. Within the demonstrator engine, an AMG will, replace the tail bearing housing radial bearing. Current engine accessories that derive power from gearbox mounted pumps will be replaced with electrical equivalents consisting of power electronically driven electrical machines, which will include fuel and oil pumps, etc. and two variable inlet guide van actuators. The engine technology platform will be based on the Rolls-Royce Trent 500 engine.

There are several key design challenges, one to overcome in embedding electrical machines within the engine is resistance to the harsh environmental conditions. The HPSG will be operating in temperatures of 300- 400°C and the FSDC in temperatures of 500-600°C. The vibration profile is equally challenging. Other considerations include the space constraint and electromagnetic compatibility. Cooling of the power electronic modules (PEM) is crucial. There is a distributed control system whereby each prototype sub-system on the demonstrator engine will form a node on a distributed control system. As far as reliability is concerned. Although the units being tested on the demonstrator are prototypes, each partner is ensuring that the issue of reliability is tackled as if the unit was for a production engine. It is concluded that a more electric engine incorporating all the technologies discussed would eliminate many systems from the engine and simplify the interface between the engine and the aircraft.

An overview of the TIMES programme was given by Ray Collins of Goodrich Power Systems. The programme has a duration of 42 months and is divided into seven work packages, as follows: system trade studies; system technologies; simulation and validation; system build and validation; final report and recommendations; programme management; and high voltage studies.

The consortium for the TIMES programme comprises the following: Goodrich Power Systems (Project Leader), Airbus UK, BAE Systems, Smiths Aerospace, FHL, Birmingham University, Cranfield University, Moir Associates, FR-HiTemp, Qinetic, Honeywell, and Rolls-Royce. The hardware demonstrator will evaluate the total electrical system characteristics, from the generation on then engine, through the electrical distribution system, to the end user loads such as electrical actuation.

Work package 2 (Simulation and Validation) has particular mention as it includes the modelling of individual system components for validation against hardware demonstrators, including electrical generation and distribution systems, flight control system, fuel system, propulsion system, complete dynamic electrical system, and the investigation of interactions between system elements. Work package 7 (High Voltage Studies) investigates the application of higher voltages (230 V VF AC, 270 V DC and higher) for future aircraft systems.

Innovative systems approach

From Graham Dodds of Airbus UK came MEA-TIMES System Architecture which also referred to the TIMES programme and describes the more electric architectures that were judged to be suitable for A320 and A330 size aircraft (known within TIMES as A320ME and A330ME). The essential electrical network provides power to those systems/ equipments that are essential for safe flight and landing in the event of a failure, such as loss of main generation. AC power can be provided to the AC busbars by a number of sources, depending upon the failure. Electrical load management will be by system located in the forward avionics bay, consisting of main distribution; high power devices located in the two primary panels, relays and circuit breakers into the secondary panels, and first and second distribution; relays and circuit breakers packaged in two identical secondary panels.

The actuators that were proposed for the MEA were electro hydraulic actuators (EHA) or electro mechanical actuators (EMAs) with the former more likely. The issue of jam tolerance on EMAs is the main reason why they have been rejected for use on primary flight control surfaces in the past. It is likely that failure conditions will provide the same level of functionality as in conventional aircraft. The likely evolution from the MEA would probably include a move away from 115 V AC CF, which has already been started. Overall, it is concluded that there would be a marginal increase in direct.

Overall, it is concluded that there would be a marginal increase in direct operating cost, but the aircraft could be flown with the knowledge that in the majority of cases, the main aircraft systems are more reliable than the conventional aircraft.

From various authors at the University of Birmingham, a status update was provided in Modelling and Simulation for the Evaluation of Electric Power Systems of Large Passenger Aircraft. This concerned the work that is being done as part of the TIMES project. The paper presents example results from the harmonic budget work and the dynamic power quality simulations. A block diagram of a typical EMA is shown. An alternative configuration is the electro hydraulic actuator where the mechanical gearbox/ ballscrew block is replaced by a bi- directional variable speed pump and hydraulic ram. The simulations in the paper are based on a single essential bus form a wide-bodied MEA similar in size to the Airbus A330.

One simulation considered involved a 12-pulse converter-based EMA aircraft system, another a DC-fed aircraft actuation system and a third a matrix converter-based EHA aircraft actuator system. The alternative systems are compared on the basis of:

1. 1.

   voltage regulation at the DC input to an individual actuator for the 12-pulse rectifier-based DC systems, and at the equivalent d-q actuator input for the matrix converter-based system; and

2. 2.

   voltage total harmonic distortion (THD) at the point of common coupling of the loads.

The DC system is shown to be the most robust architecture, being the only system to continue operating through the most severe supply transient. The results indicate the potential to educe the total capacitive stored energy in a DC system, bringing cost and weight savings, without degrading the supply transient immunity below the level of the AC systems.

Electrification of the Environmental Control System came from Jaques Herzog of Liebher Aerospace Lindenberg which noted that as the Penvironmental control system (ECS) is one of the largest power consumers among the non-propulsive systems, it is most important for this to be electrified for the MEA. Currently the ECS packs of large aircraft use air bled from the engine compressor to provide conditioned air (flow, pressure, temperature) to the cabin. Benefits of the electrification of the ECS are numerous, some being as follows.

There is no direct-intervention of the ECS in the thermodynamic cycle of the engine. Also, a high efficiency engine can only be realised with a high bypass ratio (fan flow divided by core flow) and the impact on the SFC of the engine air bleeding is higher the lower the core flow (the percentage of bleed air) related to engine core flow is higher for an engine with a high bypass ratio the core flow is reduced to a minimum. Other advantages are that an Electrical ECS will absorb the power needed to perform its tasks and not dissipate it, fuel consumption would be reduced in certain flight phases, and due to the thermal inertia of the complete ECS, it could be beneficial at aircraft level to accept a slight increase of the cabin temperature to allow other systems consuming for a short time a lot of power to be activated. Thus, there are advantages with electrification of the ECS but also some challenges, including that, although the engine may be made simpler, the ECS will be more complex.

A paper which investigates the feasibility of using Electro-thermal Ice Protection for Leading Edges on Large Aircraft was presented by various authors. The system is designed as a replacement for the current bleeds air anti-ice system. A number of electrically powered ice protection systems (IPS) have been suggested. Electro-thermal IPS present a significant power distribution and control challenge as the loads will be split into a number of zones distributed over the protected areas of the aircraft. The traditional heater elements have drawbacks and are fed from a common power bus, each with its own method of electrical isolation. An improved method of power distribution and control is required that offers reduced weight and improved reliability. Such an approach would be to use remotely located power controllers positioned close to the load requiring control.

One ice protection control system (IPCS) consists of two levels of control. There would be overall control and a number of units for local control and power switching of the heater mats. One possible heater mat layout could consist of four main sections: upper surface, lower surface, the leading edge breaker strip and the parting strips. The resistive heater mats are made from Thermion fabric. Various leading edge implementations are possible, the method depending on geometrical limitations. Thus, it is feasible to replace a bleed air IPS with an electro- thermal alternative by using new architectures and materials which is competitive in weight and reduces overall power consumption.

Power generation and distribution

A discussion of the FSDG was by several authors from Goodrich Control systems. Two drivers for change in electrical generator technology have emerged from MEA development studies. With increasing electrical power requirements the level of power drawn from the high-pressure spool is approaching the point at which engine efficiency and stability are compromised, necessitating the investigation of alternative power sources. Also, the high-reliability, lower cost future MEA will require a shift away from the gearbox-mounted three-stage wound-field synchronous generator technology used on most aircraft today.

The FSDG programme was described by Goodrich Power Systems which leads the project in partnership with other companies. It aims to demonstrate a switched reluctance generator embedded within the core of the Rolls-Royce POA engine during 2005. It will be mounted in the engine tail cone and driven by the low pressure spool (fan shaft). It is designed to produce 150 kW during normal operation up to engine red line speed and 25 kW of emergency power at engine wind-milling speed.

The FSDG will, be mounted in the engine tail cone and driven by the low pressure spool (fan shaft) via a step-up gearbox. In fact, when coupled to either the high or low pressure spool, advantages of both size and weight of some components, area realisable. Embedding one or more electrical machines with the core of the engine have a number of implications on the engine design, related to the high levels of electrical power transmitted. Owing to inaccessibility, reliability is paramount and a brushless machine format is required and cooling is very important. The bus voltage and output power requirements were chosen for the test rig and may be subject to change. The target mean time between failures (MTBF) is 40,000 h, compared with the MTBF of conventional wound field synchronous generators of 30,000 h.

The electrical and mechanical FSDG design parameters necessitate an in-line epicyclic gearbox which makes it possible to contain the dimensions of the generator with the available space envelope. Advanced techniques make reduce switching loss possible which improves converter efficiency and hence thermal performance. Trade-study work has led up to the definition of the current FSDG project, which will be further developed for testing in 2005.

From Hamilton Sundstrand came Aircraft Electrical System Architectures to Support MEA. This paper provided guidance in the process of developing such architectures and historical background to prior work in this area. Reduction of an aircraft's multiple secondary power subsystems to a single electric subsystem has been anticipated for a long time, and may be said to have started with FBW in the 1970s and 1980s. While FBW did not the aircraft's secondary power system, it did herald the digital age bringing forward the methods to design and verify the integrity of software-based systems.

Civil applications tend to be somewhat easier to categorise and standardise than military ones. A few requirements contain particular significance for MEA approaches. These include extending the failure mode and effects considerations to any electric system supported cooling system and requiring an additional (auxiliary) power source be available to support engine starting and essential services during loss of all primary (engine driven) power.

Studies have been undertaken and industry seems to have moved away from the assumption of a variable voltage/variable frequency type starter/ generator. A generic all electric sub- system architecture might be suggested that includes the following: two starter/ generators per main engine and APU and dual function engine start/ECS compressor motor drive. Also, standard 115/200 V ac depending on feeder length, equipment compatibility issues, etc.; similar primary and secondary power distribution concepts as with prior architectures; and full electric actuation with appropriate selection of actuation technologies. In addition, an air-cycle or vapour- cycle-based environmental cooling system approach is required; and an emergency generator driven by a ram air turbine or potentially, the engine's low pressure spool; as well as electric wing anti-ice and de-ice. There are numerous generation, conversion and distribution choices to be made for this architecture. Although MEA technologies are finding applications on both civil and military aircraft, their introduction has not been so rapid as once envisioned, although the benefits are being realised. Careful application of the necessary system integration and analysis tools is necessary in order to determine their optimum approach.

Actuators and effectors

Dominique van der Bossche of Airbus spoke of More Electric Control Surface Actuation of a standard for the next generation of transport aircraft. For a long time the control surfaces of transport aircraft above a certain weight have been hydraulically powered. The most recent generation of in-service commercial transports is showing generalization of the electrical signalling of the hydraulic flight control actuators, known as FBW systems. The very new aircraft generation under development will feature a mixed flight control actuation power source distribution, associating electrically powered actuators with conventional FBW hydraulic servocontrols. This paper highlights the drivers for this evolution.

The primary flight controls are dedicated to the control of the attitude on the roll, yaw and pitch axes, and the trajectory of the aircraft, with the secondary flight controls, also identified as the high lift system, dedicated to the control of the lift on the wing. The architecture of the flight control system, in terms of number of actuators per surface, number and distribution of power sources and flight control computers, is primarily driven by safety considerations. An example of the current flight control system technology is illustrated which shows the Airbus A340 distribution of actuators, hydraulic systems identified as yellow, green and blue, and flight control computers over control surfaces.

It may be asked why a challenge should be issued to the current flight control actuation power sources? It is however difficult to identify further simplification of the system on its own and major advances have to be sought elsewhere. Technology advances in electric actuation provide areas for improvement. One hydraulic system can be eliminated and replaced by a set of electrically powered actuators with no detrimental impact to the probability of losing the flight control actuation system.

The A380 is the first More Electric flight control commercial transport. The selected power source configuration features two hydraulic systems, Green and Yellow, and two electrical systems. The current Airbus FBW actuation system is based on an active/stand-by actuator arrangement. The selected More Electric architecture keeps this principle, the normally active actuators still being regular servocontrols, the stand-by actuators being changed to electrohydrostatic actuators (EHA). This is basically a self-contained hydraulic actuator incorporating a pump driven by a variable speed electric motor, by transferring the fluid from one cylinder chamber to the other, the pump and electric motor achieve the control of the position of the piston connected to the surface. The benefits of such an arrangement in terms of safety include increased power source redundancy and an increased margin of safety resulting from the introduction of the hydraulic/ electric dissimilarity in the power sources.

As regards the EHA versus the EMA, the latter is potentially more attractive in some respects although in three areas, the EHA is still preferable. These are:

1. 1.

   the jamming problem can be more easily assessed with the EHA;

2. 2.

   the same applies to the prediction of the wear life; and

3. 3.

   the introduction of EHA in parallel with regular servocontrol in the basic architecture mentioned is easier than EMA.

Further advances can be anticipated in the field of optimised utilisation of on- board power sources which will bring performance and cost benefits.

William R. Schley of Parker Aerospace spoke on the State of the Art and Remaining Challenges inn Electric Actuation Flight and Propulsion Control. The company's contributions to EHA technology are many, beginning in 1983. In the 1990s Parker demonstrated a dual-tandem EHA for one application, which was subsequently scaled-up for another one. The experience over the years enabled the company to work on the Joint Strike Fighter Integrated Subsystem Technology (J/IST). A typical J/IST actuation system architecture has each of the EHA's mechanically dual- redundant and capable of about 13hp output, but this was electronically load- rate limited to 11hp. The associated dual redundant power drive electronics (PDE) provided motor drive, velocity loop closure, load sharing, and part of the redundancy management functions. There was a separate PDE for each EHA.

The triplex control electronics (CE) unit performed position loop closure and the remainder of the redundancy management functions. The CE was shared by all five EHAs such that they could all continue to operate with any two CE channels failed. The CEs and PDEs communicated via redundant data buses. The success of EHA technology on the J/IST demonstration programme paved the way for this advance to be incorporated into the Lockheed Martin Joint Strike Fighter and this will be the first production fighter aircraft to use EHA technology for primary flight controls.

From Claverham Ltd came Electromechanical Actuation of Rotorcraft and Wide Bodied Airframes. In recent years the improvement in electromechanical technology has served to erode the clear advantage to be found within hydraulic actuation, the power density has improved and multi redundant architectures have been adopted so that actuator jamming can be safely tolerated. The inherent advantages of electromechanical actuation have also become obvious.

Redundancy within electromechanical actuators may be achieved with a variety of alternative strategies. Here, this will be limited to rotorcraft swash plate or aircraft aileron actuators. In the case of rotorcraft it is essential that full actuator authority is maintained on all control axes throughout the duration of the flight; in the context of aileron actuation, limited authority failed actuators may represent an acceptable operational mode. A second key difference is the acceptability of single point mechanical failures within aircraft applications. Generally, rotorcraft have accepted single point mechanical failures, whereas fixed wing civil aircraft do not. Finally, the application will place different constraints upon the actuator configuration.

This has led Claverham to adopt two disparate architectures in primary flight control applications, either torque summing or position summing. A large number of the challenges faced in electromechanical actuation centre around the proper management of energy within the system. This can be divided into the operational challenges of duty cycle/standing load; noise measurement accuracy; cross lane harmonisation and failure modes; and environment. The duty cycle is among the most basic of requirements for the actuator, it typically defines operational loading conditions and accelerations required throughout a series of flight events in a chronologically representative order. Noise is always present in electrical systems and any noise entering the actuator control loop at frequencies below the system bandwidth must be viewed as an integral component of the duty cycle. Because of the additional degree of freedom of the position-summed actuator, failures may be identified more readily. The environment can have a significant effect on the overall performance of the actuator and energy management. Of the two actuator technologies mentioned, Claverham will deliver the HEAT torquTe actuator for first flight trials.

The use of a Matrix Converter for Electro-Mechanical Actuation of an Aircraft Rudder came from Smiths Aerospace Mechanical Systems. A system model has been developed including the matrix converter, supply impedance, electric motor, and the various control loops and actuator loading. The model focuses on the dynamic system response.

The developed system model comprises a three-phase voltage source, supply impedance, input filter impedance, matrix converter with associated modulation algorithm, control loops, electric motor, actuator and load. The model is implemented in modular blocks. The results of the simulation illustrate the response that can be achieved for an actuator of a civil aircraft rudder with the parameters used. The dominant parameters that affect the speed of response in changing the position reference are, the maximum speed of the electrical machine, inertia of its rotor, and the maximum current. Further work will be carried out on a physical demonstrator to validate the performance in a practical implementation.