Patent Description:
Conventional thermal combustion engine driven aircraft are designed such that peak efficiency occurs during phases of peak fuel consumption such as, for example, during takeoff or climb. Accordingly, such systems are designed with engines which are oversized for less strenuous phases of flight such as cruise and descent and, thus, operate at below peak fuel efficiency for the majority of the duration of flight.

Hybrid propulsion systems for aircraft seek to improve fuel efficiency by taking into account the different operational power requirements during different phases of flight. In hybrid propulsion systems an electrical motor and thermal combustion engine are provided in series or parallel to meet the varied thrust requirements of the aircraft during different phases of flight.

In this context, an established approach is to use a so-called pulsed hybrid architecture wherein thermal combustion engines are used during takeoff and switched off during cruise in favour of an electric motor thereby avoiding operating in modes of low fuel efficiency. For example, <CIT> discloses pulsed power propulsion systems for aircraft wherein an electrical propulsor is powered by either a turbine engine or an energy storage system. During phases of flight in which the turbine engine would typically operate at less efficient throttle settings, the system shuts down the turbine and instead drives the propulsor with energy from the energy storage system so that the aircraft can be propelled without the need for the turbine to be running. The turbine may be restarted as needed, or pulsed, to recharge the energy storage system.

Similarly, <CIT> discloses a hybrid propulsion system wherein a gas turbine engine drives a propulsor via a clutch coupling in a first mode of operation, and the gas turbine engine is decoupled and an electric motor drives the propulsor in a second mode of operation. During the first mode of operation, the electric motor may be used as a generator to charge the battery.

However, such pulsed hybrid propulsion systems still require engines which are oversized for the longest phases of flight such as cruise. Though fuel efficiency is improved by only operating the engine as needed, the increased weight and drag associated with such oversized systems still result in an overall increase in fuel consumption relative to systems which are not oversized for cruise phases of flight.

In this regard parallel hybrid propulsion architectures which can cooperatively generate thrust from both thermal combustion engines and electric motors are of keen interest. In such systems thermal combustion engines can be designed, or sized, for less strenuous phases of flight such as cruise and descent. Thus, the resulting engines are not oversized for the majority of phases of flight and do not contribute to an excess in weight and drag associated with oversized engines.

In these parallel hybrid architectures the main propulsor is driven by two means; a traditional thermal combustion engine and an electric motor. The mechanical outputs of the thermal combustion engine and electric motor are combined to drive the propulsor. Typically, the thermal engine is controlled with commands from the pilot. The division of power flow from the thermal engine and electric motor can be selected through the system design and can be adjusted during operation. Typically for this type of system, the electric section is sized with sufficient energy storage to provide a power boost to the thermal engine during take- off and climb flight phases with the thermal engine then providing the full power output during cruise.

This provides a fuel saving since the thermal engine can be optimised for optimal efficiency during the cruise phase and does not need to be oversized for the full take off power.

In the state of the art parallel hybrid architecture for electric aircraft propulsion, the battery is recharged through a ground based charger after the aircraft has landed. This is similar to how a traditional non-hybrid aircraft is refuelled with jet fuel. However, the charging time for the hybrid architecture is much longer and may take several hours to fully recharge the batteries.

Moreover, parallel hybrid propulsion architectures are most applicable for relatively short flights in the region of <NUM>-<NUM> (<NUM>-<NUM> nautical miles).

To be profitable, airlines schedule short turnaround times which do not allow sufficient charging time needed to fully recharge the batteries in the hybrid system.

In addition, current technology battery systems do not have enough energy density to allow for a large amount of electrical reserve energy for use in aircraft emergency situations. In the case of an engine out or other aircraft emergencies where additional engine power is needed for extended amounts of time, the battery may become depleted and leave only thermal power available.

Recharging the batteries whilst in-flight has been contemplated. For example, <CIT> discloses a long range hybrid electric airplane wherein both a combustion engine and an electric motor drive a propeller during take-off. During cruising only a combustion engine is used and said engine operates so as to continuously generate an excess of power. This excess power is used to generate electricity for in-flight operations and to recharge batteries.

<CIT> discloses a parallel hybrid gas turbine propulsion system wherein an electric motor is connected to a shaft of the engine and configured to cause the motor to provide boost power to the gas turbine engine during takeoff. Once in a cruise mode of operation the motor is not required to provide supplemental rotation of the gas turbine engine. However, due to the physical connection rotation provided to the motor, the motor can be switched to a regenerate mode to provide electric power back to the power distribution system to power on-board electric systems and charge an energy storage component.

<CIT> discloses a propulsion system configured to reduce minor or low cycle damage of a gas turbine engine.

Despite the above, there remains a need for a system which can dynamically balance the requirement of increased fuel efficiency and in-flight recharging of an energy storage system.

In accordance with an aspect of the invention, there is provided an engine system for an aircraft according to claim <NUM>.

This engine system allows for additional in-flight electric power generation using the excess capacity of the thermal combustion engine whilst improving overall engine efficiency. More specifically, this system can increase fuel consumption within a specific operational range wherein engine efficiency increases in tandem with fuel consumption. Thus, this system allows for highly cost effective power generation which can be used to recharge the associated energy storage system (e.g., batteries), thereby reducing costs associated with extended turnaround times as batteries are charged on land. Further, the additional power stored in the recharged energy storage system can be utilised in lieu of an auxiliary power unit (APU), or to provide additional thrust in cases of emergency.

The current power output may be determined from a measured fuel consumption rate of the thermal combustion engine.

The most fuel efficient power output may correspond to a power output determined to be the most fuel efficient for the current operating conditions of the engine system.

The engine controller may further comprise a database and the optimum power output may be determined from a schedule of optimum power outputs stored on the database.

The optimum power output may be a fixed and/or predetermined value based on one or more operating conditions.

The power converter may comprise a variable torque actuator.

The system may further comprise an energy storage device, and the electric energy generated from the torque applied to the electric motor may be used to charge the energy storage device.

In accordance with an aspect of the invention, there is provided a method of operating the engine system as defined by claims <NUM>-<NUM> according to claim <NUM>.

The method may further comprise increasing the torque applied to the electric motor-generator by the power converter if the current power output is below the most fuel efficient power output, so as to increase the load on the thermal combustion engine, and by an amount required to increase the power output of the thermal combustion engine to the most fuel efficient power output.

The method may further comprise decreasing the torque applied to the electric motor-generator by the power converter if the current power output is above the most fuel efficient power output, so as to decrease the load on the thermal combustion engine, and by an amount required to decrease the power output of the thermal combustion engine to the most fuel efficient power output.

In accordance with an aspect of the disclosure, there is provided an aircraft propulsion system comprising an engine system in accordance with claims <NUM>-<NUM>.

In accordance with an aspect of the disclosure, there is provided an aircraft propulsion system comprising an engine controller configured to carry out a method in accordance with claims <NUM>-<NUM>.

Herewith will be described various embodiments of a system and method for use in an aircraft and other aerospace applications.

<FIG> shows an aircraft <NUM> comprising a propulsion system including one or more engine systems <NUM>.

<FIG> shows one of the engine systems <NUM> of <FIG>, which is configured with a parallel hybrid architecture in accordance with various embodiments of the disclosure. That is, the engine systems <NUM> each comprise a thermal combustion engine <NUM> including a combustor chamber <NUM> and may include one or more compressor sections (e.g., a low pressure compressor section <NUM>, a high pressure compressor section <NUM>), and one or more turbine sections (e.g., a high pressure turbine section <NUM> and a low pressure turbine section <NUM>). The thermal combustion engine <NUM> may comprise a rotating shaft or a first shaft <NUM> (e.g., a low pressure shaft) which may connect a low pressure compressor section <NUM> and a low pressure turbine section <NUM>. The thermal combustion engine <NUM> may comprise a second shaft <NUM> (e.g. a high pressure shaft) which may connect a high pressure compressor section <NUM> and a high pressure turbine section <NUM>. A shaft of the engine (e.g., the first shaft <NUM>) may be connected to a propulsor <NUM> via a connection means <NUM> (e.g., a gearbox). An electric motor <NUM> may be connected to the thermal combustion engine <NUM> via the connection means <NUM>. Alternatively, the electric motor <NUM> may be connected to a shaft of the engine (e.g., the first shaft <NUM>).

To achieve optimal fuel savings with a parallel hybrid architecture such as the one described herein, the thermal engine <NUM> may be sized for optimal fuel efficiency during cruise with the necessary extra power during take-off and climb provided by the electric motor <NUM>.

<FIG> shows a schematic of a system <NUM> in accordance with the disclosure, which may be referred to as a parallel hybrid propulsion system, and may correspond to one or both of the engine systems <NUM> described above with reference to <FIG>.

The system <NUM> comprises a thermal energy section <NUM> and an electric section <NUM> wherein the thermal energy section <NUM> comprises a thermal combustion engine <NUM> and a fuel source <NUM>, and the electric section <NUM> comprises an electric motor <NUM> and a power source <NUM>, <NUM>. A propulsor <NUM> is driven by two means, namely the thermal combustion engine <NUM> and the electric motor <NUM>.

The propulsor <NUM> may comprise a propeller, an unducted fan or a ducted fan. The propulsor <NUM> may comprise blades or aerofoils which may be fixed. The propulsor <NUM> may comprise blades or aerofoils which are variable, for example, blades which are rotatable about an axis.

Though a gas turbine engine is depicted in <FIG>, this figure is provided for reference only. The thermal combustion engine <NUM> of the thermal energy section <NUM> of system <NUM> may comprise a rotary engine, a reciprocating engine, a turbomachine (for example, a turboprop, a turbofan, a turboshaft or a turbojet engine) or any other aircraft engine as is known in the art.

The mechanical outputs of the thermal combustion engine <NUM> and electric motor <NUM> may be combined through a connection means <NUM> to drive the propulsor <NUM>. The connection means <NUM> may comprise a gear box and/or a suitable transmission (e.g. a freewheel or overrunning clutch) configured transmit drive from the thermal energy section <NUM> and the electric section <NUM> to the propulsor <NUM>.

The thermal combustion engine <NUM> may be controlled with a controller <NUM> (e.g., an engine controller), which may receive commands from an operator (e.g., a pilot), current operational conditions and/or inputs from a control system of the aircraft wherein the operating conditions includes operation state of aircraft, e.g., as set by pilot (take-off, climb, cruise etc.), operational parameters of engine (pressure, temperature, measured specific fuel consumption etc.), sensor data from aircraft (altitude, yaw/pitch/roll, outside air temperature etc.). The controller acts to produce the necessary actuation signals to drive the thermal combustion engine <NUM> (e.g., the controller may control a supply of fuel from fuel source <NUM> to the thermal combustion engine <NUM>).

The electric motor <NUM> may be controlled with the power source. The power source may comprise a power converter <NUM> which converts electrical power from an energy storage system <NUM> (e.g., a battery) into a form that can drive the electric motor <NUM>. This conversion may be from DC (e.g., from a battery) to multi-phase AC (to drive the electric motor). The output from the power converter <NUM> may be controlled by the controller <NUM> such that the electrical motor <NUM> drives the mechanical output as required by the system <NUM>.

Alternatively, as depicted in <FIG> (where like elements are referred to with like reference numerals), the connection means <NUM> may be omitted and the electric motor <NUM> may connect to a shaft of the thermal combustion engine <NUM> (e.g., in the case of a gas turbine engine, the first or second shaft <NUM>, <NUM> of the thermal combustion engine <NUM> depicted in <FIG>). The controller <NUM> and power converter <NUM> acts to control the system <NUM> as described above.

Referring to <FIG>, the division of power flow between the thermal section <NUM> and electric section <NUM> can be selected through the system <NUM> design and can be adjusted during operation using the controller <NUM>. The electric section <NUM> is sized with sufficient energy storage to be configured in a first mode of operation to provide a power boost to the thermal combustion engine <NUM> during take-off and climb flight phases. The thermal combustion engine <NUM> can then provide the full power output during cruise if desired.

<FIG> shows typical fuel efficiency characteristics over a varying output power range of a thermal combustion engine in accordance with the disclosure. A peak in fuel efficiency is achieved at a relatively high power output P2. Above and below this power output the fuel efficiency of the thermal combustion engine <NUM> is reduced. Thus this power output is considered the optimum (i.e. most fuel efficient) power output P2 of the thermal combustion engine <NUM>. During the takeoff phase the thermal combustion engine <NUM> operates to maximise power output, and typically operates at or above the optimum power output P2 (e.g., in the region of P2 or above). Additional power output or drive can be provided by the electric motor <NUM> as required by the system <NUM>.

During the cruise phase, the required power output may vary due to changing flight conditions. The required power output may vary over a range PΔ between a maximum power output required during cruise P3, and a minimum power output required during cruise P1. The optimum or most fuel efficient power output P2 resides in this range. As the power or thrust requirement from the system fluctuates (e.g., due to changing flight conditions), the required power output of the system may drop below the optimum power output P2.

Consequently, when the required power output from the thermal combustion engine <NUM> is less than the optimum power output P2 (in the range P1 to P2) then there is an additional highly fuel efficient power capacity available within the system <NUM>. Additional fuel use in this region will increase the overall fuel efficiency of the thermal combustion engine <NUM> towards peak fuel efficiency. Thus, by increasing the load on the thermal combustion engine <NUM>, additional power can be generated in a highly fuel efficient manner.

In accordance with the disclosure a method for efficiently generating power from a parallel hybrid propulsion system <NUM> as described above is now disclosed.

When the required thrust from the system <NUM> is reduced such that thrust is no longer required from the electric motor <NUM> (i.e. the thermal combustion engine <NUM> alone can provide sufficient thrust), the electric section <NUM> may be switched to a second regenerative mode of operation. Switching to the regenerative mode may occur automatically as a consequence of a reduced thrust requirement or due to other inputs to the system <NUM> (e.g., pilot input).

In the second mode of operation the controller <NUM> and the power converter <NUM> may be configured to convert the electric motor <NUM> into a generator. Thus the electric motor <NUM> may also be referred to as the generator <NUM>. In order to generate power, torque may be applied to a shaft of the electric motor <NUM>. Torque may be applied with a generator current controller such as a torque actuator which may be a variable torque actuator. In the absence of any applied torque the electric motor <NUM> may be free spinning.

<FIG> shows a schematic for a method <NUM> for generating power (e.g., using system <NUM>) comprising steps as outlined below.

At step <NUM> a current power output from the thermal combustion engine 'P' is determined. The current power output P of the thermal combustion engine <NUM> may be determined in any manner as is known in the art. For example, power output may be determined from specific fuel consumption as is measured by a fuel pump or flow meter or the like. Power output may be determined with sensor data from the thermal combustion engine <NUM> (e.g., engine torque and speed measurements from suitable sensors), the connection means <NUM> and/or the propulsor <NUM>. Power output from the thermal combustion engine <NUM> may be determined with data collected from or associated with the electric motor <NUM> (e.g., rotational speed from suitable sensors).

At step <NUM>, the measured or determined power output P is then compared with the optimum power output P2. The optimum power output P2 may vary dynamically with operating conditions, and thus may not be constant. A schedule for the optimum power output P2 may be determined by design and testing of the thermal combustion engine <NUM> or could be measured dynamically through measured fuel usage at or from the fuel source <NUM>. The optimum power output P2 may be determined or scheduled in relation to different operating conditions, wherein the operating conditions includes operation state of aircraft, e.g., as set by pilot (take-off, climb, cruise etc.), operational parameters of engine (pressure, temperature, measured specific fuel consumption etc.), sensor data from aircraft (altitude, yaw/pitch/roll, outside air temperature etc.).

If the current power output P is found to be below the optimum power output P2, then the system may be configured to increase the torque on the electric motor/generator <NUM> (step <NUM>), in order to increase the load on the thermal combustion engine <NUM>, thereby increasing fuel consumption to a more efficient or optimum level. The additional power generated may be used to charge the energy storage system <NUM>.

Conversely, if the current power output P is found to be above the optimum power output P2, then the system may decrease the torque on the electric motor/generator <NUM> (step <NUM>), in order to decrease the load on the thermal combustion engine <NUM>, thereby decreasing the fuel consumption to a more efficient or optimum level.

Increasing the power output to charge the energy storage system <NUM> in this way allows the thermal combustion engine <NUM> to operate at a higher or peak fuel efficiency. This means that the energy storage system <NUM> can be charged in-flight, reducing (or even eliminating) the time taken to recharge on the ground. This may also provide additional reserve power for use in emergency conditions.

Referring again to <FIG>, the controller <NUM> may be configured to carry out any or all of the above method steps. A communications data link may be used to provide coordination between the controller <NUM> and the electric motor/generator <NUM>. For example, the controller <NUM> may be configured to identify operating conditions where additional capacity within the PΔ range is available, and then output a signal to the power converter <NUM> to start to extract additional power by applying torque with the power converter <NUM>. This power may then be delivered to the energy storage system <NUM>.

The above described method may be performed with a control system as shown schematically in <FIG>, which may be referred to as a closed loop control system. The controller <NUM> may comprise a summation unit <NUM> and a proportional integral differential (PID) controller <NUM>. The summation unit <NUM> may generate an error signal by comparing the current power output P with the optimum power output P2 of the thermal combustion engine <NUM>. As described above, the optimum power output may vary dynamically. Thus, the controller may further comprise a database <NUM> containing a schedule of optimum power outputs for various operating conditions. Alternatively, the controller may calculate the optimum power output P2 dynamically through measured fuel usage at or from the fuel source <NUM>, or the optimum power output P2 may be a fixed value corresponding to a set rate of fuel consumption.

The error signal from the summation unit <NUM> may be calculated by subtracting the measured thermal combustion engine power output P from the optimal power output P2. The error signal is passed through the PID controller <NUM>, which is configured to generate a generator power demand signal <NUM>, which is passed over the data link to the power converter <NUM>. The generator power demand signal <NUM> corresponds to the additional power which can be generated in order to reach peak fuel efficiency. The power converter may comprise means <NUM> for determining a generator speed and the power converter <NUM> may comprise a controller <NUM> which is configured to control a current generated by the generator <NUM> (e.g., a torque actuator).

The power converter <NUM> may comprise means (e.g., circuitry) <NUM> configured to divide the generator power demand signal <NUM> by the generator speed <NUM> to generate a torque command <NUM>. Alternatively, such calculations may be conducted by the controller <NUM> alone. The torque command <NUM> may correspond to the actual torque applied on the generator <NUM>, or it may correspond to a differential in the amount of torque applied (i.e. the change in torque). The torque command <NUM> may then be applied as an input into the current controller <NUM> which increases the torque on the generator <NUM>, thereby increasing the load on the thermal combustion engine <NUM>. This increases the current power output P and fuel consumption to a more efficient state.

Conversely, if the current power output P is greater than the optimum power output P2, then the generator power demand <NUM> may be negative. This would cause the torque command <NUM> to be negative and would result in a reduction of applied torque, or complete disengagement of the current controller <NUM> from the generator <NUM>, thereby reducing the load on the thermal combustion engine <NUM>.

Claim 1:
An engine system for an aircraft comprising:
a propulsor (<NUM>) configured to drive an aircraft (<NUM>);
a thermal combustion engine (<NUM>) configured to drive the propulsor (<NUM>);
an electric motor-generator (<NUM>) connected to the thermal combustion engine (<NUM>) and configured to drive the propulsor (<NUM>);
a power converter (<NUM>) configured to apply a torque to the electric motor-generator (<NUM>) and generate electric energy from the torque; and
an engine controller (<NUM>) configured to:
determine a current power output (P) of the thermal combustion engine (<NUM>);
determine a most fuel efficient power output (P2) of the thermal combustion engine (<NUM>) based on current operating conditions;
compare the current power output (P) with the most fuel efficient power output (P2); and
increase or decrease the torque applied to the electric motor-generator (<NUM>) so as to increase or decrease a load on the thermal combustion engine (<NUM>), wherein the torque is increased or decreased by an amount required to increase or decrease the power output of the thermal combustion engine (<NUM>) to the determined most fuel efficient power output (P2);
wherein:
the engine controller (<NUM>) is configured to increase the torque applied to the electric motor-generator (<NUM>) by the power converter (<NUM>) if the current power output (P) is below the most fuel efficient power output (P2), so as to increase the load on the thermal combustion engine (<NUM>); and/or
the engine controller (<NUM>) is configured to decrease the torque applied to the electric motor-generator (<NUM>) by the power converter (<NUM>) if the current power output (P) is above the most fuel efficient power output (P2), so as to decrease the load on the thermal combustion engine (<NUM>).