It is theorized that systems and equipment for an All Electric Aircraft (AEA) will be developed 20 years hence as a More-Electric Aircraft (MEA) with no bleed system, which is a concept typified by the Boeing 787, and electric powered propulsion (including electric distributed thrust or electric hybrids by gas turbine power generation), which is expected to be realized after the 2040s. In this trend, a More-Electric Engine (MEE) plays the roles shown in Table 1.

A conventional engine has two responsibilities: propulsion as a primary function, and a power plant for the secondary systems power source as a subsidiary function. This article will focus on the expectation that MEE will be provided as a future key technology for the total energy management of MEA. Specifically, the engine will be an important factor not only for the conventional role of high efficiency and low emission, but also for optimization between the propulsion system and electric power system because an MEA integrates power management into the electric power generated by engines. In addition, the MEE will also have to consider the importance for optimization of the aircraft in integrated management by advancing management of fuel burns and the thermal control system, optimization of engine control and information sharing with the entire aircraft, and the integration of flight control.

Low-Pressure Spool Power Generation

Table 1. Roles of the MEE of electric aircraft
The conventional fuel control system wastes a lot of power because a fuel pump driven by engine extraction power through an accessory gear box delivers excessive fuel flow compared to the necessary fuel flow to acquire a specified thrust, which circulates in the bypass circuit of fuel metering. An electric fuel system that eliminates excess flow improves the unnecessary rise of fuel temperature, and the fuel for a heat sink is expected to be a solution for the exhaust heat problem of aircraft. Since the electric fuel system affects the engine thrust as the discharge changes, a fault tolerant design in which the discharge does not change during a fault is necessary.

A similar design concept must also be considered for the power supply of the electric system. Although a starter/generator realized by the high power of the aircraft's generator contributes to the reduction of the conventional air starter and plumbing, it is considered a transfer mechanism that may become larger due to increased extraction of power by a power take-off as the secondary power, which was bled of air, replaced with electrical power by an engine-driven generator, disposal of the generated heat by the enlarged starter/generator, and the limitation of location and available space around an engine.

Electric Power Demand and Engine Control

Table 2. Merits and demerits of the extracting power shaft of the engine
No-bleed systems of the MEA generate more electric power because they transfer the secondary power by compressed air from the conventional compressor to power generation by the rotational energy of a high-pressure spool shaft. The AEA, with a single aisle and 150 seats, may need 1MW of electric power, or for an eight-hour flight, 7.8MWh. In order to realize these electric power demands, the engine system plays a role in power generation in the total energy management system that consists of power generation, distribution, storage and consumption.

The power generation system is necessary for the engine system considering the improvement of generating efficiency for future increased power demand, stability of engine control, and limits of mounting structures of an extraction mechanism and generators. General aero-engines have a low-pressure (LP) spool and high-pressure (HP) spool in the compressor and the turbine, respectively, and fans that acquire thrust are driven by a low-pressure shaft and the generator and the pumps are driven by the torque of a high-pressure shaft. The generator drive by the torque of a low-pressure shaft should be examined in consideration of weighing up the merits and demerits of the high-/low-pressure shafts shown in Table 2.

Integrated Propulsion Control

Table 3. Research targets of the integrated propulsion control
The electric power management needs to be linked with every status and situation of the aircraft and requires a stable power supply and quick response in an emergency. A conventional fuel pump driven by mechanical power needs to be mounted on the engine AGB, but for the electric power system it needs to provide flexibility upon being mounted on the fuel pump, and fuel system integration should be assumed from the point of a holistic and optimized system.

Tightening the information and communication between the control system of MEE and the aircraft system may enable integrated control such as removing limiting conditions temporarily by analysis of the information and the status in coordination with engine control and aircraft control in an emergency. Coexistence of aircraft control, electric power management and stability of engine operation must function as total energy management at normal times and provide safety and stability in an emergency (Table 3).

As described, the MEE shares a close working relationship with the MEA and the AEA in every aspect. While the MEE, especially the electric fuel system, is thought to play an important role in thermal management and control of electric aircraft in the future, it is also an essential system as well as a flight control system and requires high reliability and safety.

Sufficient Emergency Power

Figure 1. Proposed power bus architecture for AEA
For the AEA, including an electric fuel control system, the electric power system is one of the crucial systems and requires high stability and fault tolerance that do not cause any power fluctuations or disruptions by a system fault. The authors propose electric bus designs such as a power interchange through the MEAAP (More-Electric Architecture for Aircraft and Propulsion) research. However, our targeting emergency power source capacity was merely assumed as a substitute of the power source at normal times and has not mentioned the observation that the AEA is feasible under circumstances such as generator failure and a disabled bus, or while the engine is off. Therefore, a backup or emergency power source should be large and complicated.

For example, the ram air turbine (RAT) system has covered essential power supply and the flight control drive for conventional aircraft in an emergency, but it may not be enough for the AEA due to increased power demand. The auxiliary power unit (APU) is also not sufficient to back up the AEA because it has a failure mode as common as the engines, such as a disabled fuel system. Therefore, alternative energy sources such as batteries and fuel cells (FC) will be required. In that case, the problem is specific energy performance and it is difficult to mount those batteries and fuel cells due to the weight. Taking this into consideration, we’ve focused on an LP power generation system as a high-performance emergency source.

Definitions of Power System Operation

Electric Power Demand for Emergencies

There are two modes of power supply system operation: ground operations and in-flight operations.

Table 4. Considerations for the necessary electric power of a single aisle aircraft with 180 seats.
For in-flight operation, generally, the volume of an engine generator is considered to be set to satisfy the total electric energy demand for the necessary equipment. Therefore, the integration of the total electric energy is calculated. As for ground operations, basically, the engine generator provides power while the engine is turned over even if it is in a ground idle condition. However, the secondary power supply (for air conditioning, starting avionics, galley, etc.) is necessary for ground operation before the engine starts and generally electric power and compressed air are provided by the GPU or the APU. When the engine starts, the air starter requires compressed air or the starter/ generator requires electric power. A power supply source for them is required and the APU dedicates starter driving power to conventional aircraft. This power supply and demand needs to be met in order to consider the alternative function of the APU. Those would be considered normal operations.

In terms of emergency and recovery operations, first, it should be determined whether at least one or more engines work correctly, or if both engines of a twin-engine aircraft are down in an emergency situation. If one or more engines work correctly, the capacity of the engine generators are thought to be designed to basically enable normal operation. In these cases, the following need to be satisfied from the requirements of each certification.

The Control System can be Operated and the Aircraft Can Fly and Land Safely Even if Two Engines are Broken
Applicable requirement;
CFR(Code of Federal Regulation) Title 14, Part25,
Subpart D - Design and Construction, CONTROL SYSTEMS,
§ 25.671 General. (d)

In this condition, electric power must be provided to the avionics, the flight control, a high lift device so that communication, position and attitude, attitude control and lighting is complete, but not to the ECS, the galley, the fuel pumps (the engines are already down), on-board entertainment, anti-icing systems (it was impossible for conventional aircraft to provide the electric power while the engines were down because bleed air was used).

The Engines Can Restart at a High Altitude Even if the Normal Power System Cannot be Used
Applicable requirement; CFR Title 14, Part25,
Subpart D, ELECTRICAL SYSTEMS AND EQUIPMENT,
§ 25.1351 General. (d)

The starter/generator needs to be used for engine restart, which consumes a lot of electric power. Since the time to restart the engines is short, electric power is provided only to the communication system, the attitude indicator, the attitude control and the fuel pumps but not to all other nonessential and non-urgent-use equipment. After restarting the engines, all the systems are restored immediately.

The ETOPS Should be Satisfied Even if a Triple Fault in the Electric Power Equipment Occurs
Applicable requirement; CFR Title 14, Part25, Subpart I - Special Federal Aviation Regulations, APPENDIX K TO PART 25-EXTENDED OPERATIONS (ETOPS),
K25.1 Design requirements.
K25.1.3 Airplane systems. (b)
K25.1.4 Propulsion systems. (a)

Electric power is supplied to the communication system, the attitude indicator, the attitude control system, the lighting, the fuel pumps and anti-icing systems so that normal operation is available with one engine. Table 4 shows the calculated result of the necessary electric power of a single aisle aircraft with 180 seats based on the above conditions. Power demand for each electrical load was estimated assuming the 180-seater AEA system.

The Peak Cut Operation

It is necessary to reduce the electricity in an emergency and the electric power for the flight control is estimated at approximately less than 30% of the maximum electric power by the foregoing calculations. For the AEA, all actuators for flight control are electric and the electric power system is required to provide peak current and absorb regenerated current repeating powering and regeneration according to the control surface operation. Especially, when the peak current is provided from the power supply, the capabilities of the power system affect not only the generator capacity and size but also the wire capacity and weight to prevent a voltage drop by line impedance.

Table 5. Electrical power system distribution
As clarified by the typical current waveform, there is a big difference between the maximum current, which is consumed during acceleration and at maximum speed, and the average current. It means that the challenges regarding the increased capacity of the generator and line impedance are solved for the entire power system by leveling the electric power with an accumulator and controlling the maximum current. We have proposed electric accumulation and electric peak-cut control using a flywheel battery (FWB) as a countermeasure against the above. A capacitor, which is superior in the following points to an electro-chemical battery, such as a lithium-ion battery, is generally applied to the electric accumulator because: (a) it has a long life (semi-permanent repetition), (b) it can be charged by a large current, (electricity can also be discharged by a large current), (c) it is an eco-friendly material and (d) it is able to accurately comprehend the residual energy from terminal voltage.

A capacitor, which saves electric charge in the inductor, is generally applied for the electric accumulator; however, FWB, which saves the power by kinetic energy, is superior for an application requiring both high power and energy. FWB is also highly effective by being located close to the load whenever possible and therefore, superior to the capacitor, which is inferior in low temperature characteristics, considering its mounting on a non-pressurized area. In addition, it is considered that both the leveling and high reliability of the electricity can be achieved by making use of the feature that FWB can transmit the electricity and energy insulating electrically through kinetic energy and configuring to distribute the electricity to multiple control surface actuators from the multiplexed power system.

System Configuration

This system is comprised of the combination of the conventional HP-SPOOL power generation and LP-SPOOL power generation. LP power generation is available to provide more electric power than the HP power generation at the time of the engine halting, which means a windmill condition in flight. It is expected that LP power generation can be used as an alternative power generation system without mounting the RAT by adding the same function as RAT to it. The FC (Fuel cell) is a part of the configuration as an alternative system to the APU, but a large weight disadvantage is caused when the FC is requested to generate the same volume of power as the APU. Therefore, the important point of this study is to discuss how much the output of the FC can be reduced with the combination of LP power generation assuming the operational situation.

System Distribution

The system distribution of a single aisle aircraft with 180 seats is described. We assume 400kw electric power generation per engine through the combination of two HP-SPOOL power generations of 150kw × 2 and one LP power generation of 100kw, considering the function redundancy of the starter for the conventional HP-SPOOL power generation. Confirming the conformity by each condition, given that a windmill power generation of the LP generator is 1/2 the rated capacity (assumed value), it is found that mounting an FC rated about 80kw enables operation.

According to the current FC system for aviation, the weight is about 200kg per 25kw. Therefore, approximately 600kg- 800kg composition is assumed. Given that the RAT weighs 200kg including the structure and the APU weighs 200kg (including fuel), the weight is increased by about 200kg-400kg. Since, on the ground, the engines need to be started only with the power from the FC, the FC must tolerate about twice the rating for power supply for a short time for engine starting, which should be taken into consideration for the thermal design of fuel cells.

There are several approaches to developing this power source. The existing high-power ratio HEV lithium-ion battery (LIB) is 25 kg at a 30 kW rate (1.4Wh). If we assume a power source capable of 150 kW, the LIB system weight would be 125 kg. This LIB would also incorporate the engine starter current design specifications. However, this LIB system limits the time to operate the engine starter. The LIB system stores the 7.0Wh at 150kW output power and approximately 170-second run times.

Challenges of Generating Power by a Low-Pressure Shaft

The capacity of LP power generation affects the feasibility of the FC.The challenges of generating power with the low-pressure shaft as opposed to the more common high-pressure shaft, are:

  1. Type of generator
  2. Location of the generator
  3. Proposal for a new aircraft electric power system

The specific difference between a generator driven by a low-pressure spool and that driven by a high-pressure spool is rotation speed range. The rotation speed of a high-pressure spool for general aero-engines changes within the range of 60-100%, but that of a low-pressure spool changes within the range of 20-100%. For the generator of the highpressure shaft, a field winding-type synchronous type is broadly used, but it is not suitable for power generation by a wide rotation speed range because a built-in permanent magnet generator is used as a field power supply and output voltage is in proportion to the rotation speed cubed in principle.

Table 6. The features and types of generators
The PM (Permanent Magnet motor) type and the SR (Switched Reluctance motor) type are considered to be prospects for the generator of the low-pressure shaft. The features of these types are shown in Table 6. The PM type has advantages in efficiency and output density (weight/ volume), but it has a drawback relating to fault tolerance in which it cannot stop excitation when it breaks down. The SR type is comprehensively potent in terms of fault tolerance, although it has drawbacks in weight and constitution.

The current generator of the high-pressure spool is mounted on an AGB (Accessory Gear Box). The generator of the low-pressure spool is considered to be embedded and directly linked to the engine shaft in an aft-sump space or a front-sump space, in addition to mounting on the AGB as well as that of the high-pressure shaft, and considering the reduced mounting space around a fan due to a high bypass ratio and AGB-less for any future plan. The aft-sump area (tail cone) has a large mounting space but has a relatively high temperature. On the other hand, the front-sump area is limited in mounting space. The method to mount on the AGB is thought to be an extension of the conventional technique and has fewer challenges, but the increased weight by larger PTO (Power Take-Off) and the influence on engine performance by the decreased area for air flow in the engine need to be examined.

The maximum rotation speed of the low-pressure spool is about 1/3 that of the high-pressure spool. Since the constitution of generators become larger as rated speed becomes lower if they have the same output/method, the generator of the low-pressure spool will become larger than that of the high-pressure spool. Although it can be smaller by adding a gear and accelerating, the acceleration is limited by the magnetic saturation characteristic of the magnetic body and mechanical strength. In addition, since the efficiency of the generator becomes lower, the fluctuation range of the efficiency by the rotation speed becomes large for the generator of the low-pressure shaft that has a wide range of rotation speeds. Since the generator of the low-pressure spool is inferior to that of the high-pressure spool in performance and efficiency, it is practical to think this for the relatively low power. For a test calculation, the volume is set based on an assumption that a 100kw generator is realistic.

Future Challenges

Downsizing and lightening is the most challenging matter and downsizing of power electronics using a fuel cell and wide band gap elements is an essential technique. Downsizing and lightening of generators is also important. For the future, since lightening is important to realize LP power generation, research is expected to move toward an electric motor with high temperature tolerance that can operate under temperatures of more than 400°C for LP power generation, aiming at a composition to mount the LP power generator on an AFT-sump that does not require a PTO shaft.

This article is based on SAE Technical paper2015-01-2408 by Hitoshi Oyori, IHI Aerospace Co. Ltd., and Noriko Morioka and Tsuyoshi Fukuda, IHI Corporation, doi:10.4271/2015-01-2408.