Patent Description:
In aerospace, there is a trend towards aircraft and their propulsion systems becoming 'more electric' in their design. For example, aircraft and engine systems and accessories which have previously been mechanically, pneumatically or hydraulically powered may be replaced with electrically powered equivalents. Spool-coupled electric machines with high power ratings may be utilized to increase the power generation capability to meet the increased electrical power demand. Spool-coupled electric machines may also be used to accelerate or decelerate the spools of the engines to improve engine operability, for example to improve engine surge margin and reduce fuel consumption. Hybrid electric aircraft similarly utilize a large amount of electrical power to power propulsive and non-propulsive electrical loads.

The increased use of electrical power in these engines and aircraft creates opportunities in terms of weight reduction, reduced mechanical losses and reduced fuel consumption. However, it also creates new fault modes.

United States Patent Application Publication <CIT> relates to aircraft systems and methods for powering one or more critical loads where one or more engines fail. The system includes a generator, coupled to a main engine. In case the generator fails, an auxiliary power (APU) unit is provided. Where the APU fails or is incapable of providing sufficient power to loads, an emergency power system is configured to supply power to certain loads.

European Patent Publication <CIT> relates to a vehicle comprising an engine restart system.

United States Patent Application Publication <CIT> relates to an aircraft power management system.

According to a first aspect, there is provided a method of operating an aircraft power and propulsion system. The power and propulsion system comprises: one or more propulsive gas turbine engines, each engine comprising a plurality of spools including a first spool and a second spool, a first electric machine mechanically coupled with the first spool and a second electric machine mechanically coupled with the second spool, wherein each of the first and second electric machines of each of the one or more gas turbine engines comprise a first sub-machine and a second sub-machine; and an electrical system connected with the electric machines and comprising an electrical energy storage system. The method comprises: operating one or more of the electric machines of the one or more gas turbine engines as generators to extract, for each of the one or more gas turbine engines, greater than <NUM>% of a combined mechanical power from one or more of the spools and to generate electrical power therefrom; meeting an electrical power demand, PD, of a plurality of electrical loads connected with the electrical system by supplying the plurality of electrical loads with electrical power generated by the one or more electric machines; and determining a condition to the effect that there is a fault in one or more of the electric machines and that an amount of electrical power being generated by the power and propulsion system (<NUM>) has reduced to a lower level, Pfault. The method further comprises, responsive to the determination: during a time period ΔT, meeting at least part of the electrical power demand of the plurality of electrical loads by discharging the electrical energy storage system; and during the time period ΔT, controlling the plurality of electrical loads to reduce the electrical power demand.

Thus, in the event of a fault or failure which results in a sudden drop in the power generation capability of the system, the energy storage system (ESS) automatically discharges so as to provide a time period, ΔT, in which the system power demand can be managed and brought to a sustainable level. This reduces the risk that the reduction in the power generation capability will escalate, for example due to a sudden erosion of engine surge margin that leads to a further drop in the amount of power than be extracted from the engine(s).

This approach may be particularly effective in platforms which feature a high level of electrical loading relative to the amount power produced by the prime mover(s), and/or a high level of criticality in the electrical loading. For example, in such a platform, a sudden drop in the power generation capability could result in the power demand exceeding the power generation capability, which could lead to critical loads losing power. This could itself lead to a further, possibly unrecoverable, drop in power generation capability. As another example, meeting the electrical power demand following a sudden drop in power generation capability could result in a sudden erosion of engine surge margin. If an engine surges then the power generation capability could drop further, and this could then lead to critical loads losing power.

Thus, the plurality of electrical loads may include one or more critical electrical loads. The term "critical electrical load" will be understood by those skilled in the art to refer an electrical load whose function is essential to the operation of the overall power and propulsion system or the aircraft as a whole. One example of a critical electric load is an electrically-powered fuel pump, which supplies fuel to the combustion equipment of a gas turbine engine. If such a fuel pump loses power then the combustion equipment is not supplied with fuel and will shut down, leading to a drop in power generation capability. Another example of a critical electric load is an electrically-powered oil pump which supplies lubrication to the bearings of an engine shaft or power gearbox, for example. If such an oil pump loses power then the engine could seize, leading to a drop in power generation capability. Other examples of critical electrical loads may include engine controllers (e.g. FADECs) and actuators for aircraft control surfaces.

The amount and fraction of the power extracted from the engine spool(s) by the electric machines to power the electrical loads may be higher than in conventional systems. For example, in some embodiments, prior to the determination of a fault or failure, a peak value of the percentage of the total combined shaft power of the plurality of spools of an engine extracted from the shafts and converted to electrical power is greater than <NUM>%. In some embodiments, the percentage may even be greater than <NUM>%, <NUM>% or even greater than <NUM>%. Generally, the peak percentage will be less than <NUM>%, and is preferably less than <NUM>%. For example, the peak value may be in the range <NUM>-<NUM>%. In some embodiments, a ratio defined as: a combined rated power of all of the one or more electrical machines mechanically coupled with the spools of the gas turbine engine, divided by the maximum rated thrust of the gas turbine engine, may be greater than <NUM> WN-<NUM>. The thrust rating is a dry thrust rating (i.e. without the use of reheat/afterburner, if such a system is present in the engine). The ratio may be greater than <NUM> WN-<NUM>, greater than <NUM> WN-<NUM>, greater than <NUM> WN-<NUM>, greater <NUM> WN-<NUM>, or even greater than <NUM> WN-<NUM>. The ratio may be less than <NUM> WN-<NUM>. In one group of embodiments, the ratio is between <NUM> WN-<NUM> and <NUM> WN-<NUM>, preferably between <NUM> WN-<NUM> and <NUM> WN-<NUM>. In an embodiment, the ratio is between <NUM> WN-<NUM> and <NUM> WN-<NUM>.

The time period ΔT is preferably at least <NUM> seconds, and may be longer, for example at least <NUM> seconds, at least <NUM> seconds or at least <NUM> minute.

In order to supply the required power for the time period ΔT, the energy storage system may have a high energy storage capacity relative to the platform power. For example, a ratio defined as: a total energy storage capacity of the electrical energy storage system, divided by a combined maximum rated thrust of the one or more gas turbine engines, may be greater than or equal to <NUM> Watt hours per Newton (WhN-<NUM>). The ratio may be between <NUM> and <NUM> WhN-<NUM>. In an embodiment the ratio is between <NUM> and <NUM> WhN-<NUM>. In one group of embodiments the ratio is between <NUM> and <NUM> WhN-<NUM>. In another group of embodiments the ratio is between <NUM> and <NUM> WhN-<NUM>.

The method may further comprise: responsive to the determination, during the time period ΔT, controlling the one or more gas turbine engines to adjust operating points thereof and thereby increase the electrical power generation capacity of the power and propulsion system. For example, during the time period ΔT, the rate that fuel is supplied to an engine may be increased to move the engine to a higher thrust operating point. In this way, more electrical power can be extracted from the engine without an unacceptable erosion of its surge margin. In another example, during the time period ΔT, a speed of the aircraft may be reduced. Electrical power requirements have been found to generally increase with aircraft speed, such that a decrease in speed reduces the electrical loading. The aircraft speed may be reduced by reducing the rate that fuel is supplied to an engine to move the engine to a lower thrust operating point, and/or by operating one or control surfaces of the aircraft.

Controlling the plurality of electrical loads to reduce the electrical power demand may comprise implementing a pre-defined limp mode having a reduced electrical power demand. Those skilled in the art will understand that a 'limp mode' is a pre-defined set of system operating parameters in which critical systems are operable but the overall power consumption of the system's electrical loads is reduced. Utilizing a pre-defined limp mode allows the power and propulsion to be quickly transitioned into a sustainable operating state in the available time period ΔT. A plurality of different limp modes may be pre-defined, each corresponding to one or more particular types of fault. A limp mode may then be selected from amongst the plurality of pre-defined limp modes based on a type of fault that is detected.

Implementing the pre-defined limp mode may further comprise controlling the one or more gas turbine engines to adjust operating points thereof. In other words, the limp mode may also define one or more operating parameters or settings of the gas turbine engine(s). For example, the limp mode may define a thrust setting or range of thrust settings that can be used.

If Pfault is lower than the electrical power demand, PD, of the plurality of electrical loads (i.e. PD >P fault), controlling the plurality of electrical loads may comprise: during the time period ΔT, reducing the total electrical power demand to below the reduced amount of electrical power, Pfault. Reducing the total electrical power demand may comprise controlling an operation of one or more loads to reduce their power demands. Additionally or alternatively, reducing the total electrical power demand may comprise decelerating the aircraft. Meeting at least part of the electrical power demand of the plurality of electrical loads by discharging the electrical energy storage system may comprise, during the time period ΔT, supplying a deficit in the generated electrical power (e.g. PD - Pfault) from the electrical energy storage system. Thus, all electrical loads - including critical electrical loads - remain powered until after the total power demand, PD, can be reduced to a sustainable level.

If Pfault, is greater than the electrical power demand, PD, of the plurality of electrical loads (i.e. Pfault > PD), meeting at least part of the electrical power demand of the plurality of electrical loads by discharging the electrical energy storage system may comprise: during the time period ΔT, supplying at least part of the electrical power demand from the electrical energy storage system to maintain surge margins of the one or more gas turbine engines until the electrical power demand has been reduced. Thus, even though the power and propulsion system may still be capable of providing power to all of the electrical loads, the electrical energy storage system can be used to soften the impact of the fault on the surge margin(s) of the engine(s) and/or limit peak turbine tem peratures.

The method may further comprise, prior to the determination to the effect that there is a fault in one or more of the electric machines:
maintaining a state of charge of the electrical energy storage system above a pre-defined level. This may ensure the ESS is always in a state of readiness should there be a fault that results in a sudden drop in power generation capability. The pre-defined level may be sufficient to power one or more critical electric loads (e.g. an electric fuel pump) during the time period ΔT.

The method may further comprise, prior to the determination to the effect that there is a fault in one or more of the electric machines:
within a pre-defined time after take-off of the aircraft, charging the electrical energy storage system to a state of charge greater than or equal to a pre-defined level using electrical power generated by the one or more electric machines. The ESS may be utilized during take-off or climb, for example to add power to the engine spools to reduce fuel consumption or noise during these flight stages. After this, the ESS may be recharged to at least the pre-defined level so that the ESS is in a state of readiness should there be a fault that results in a sudden drop in power generation capability. The pre-defined level may be sufficient to power one or more critical electric loads (e.g. an electric fuel pump) during the time period ΔT.

The energy storage system may take the form of one or more batteries and/or one or more supercapacitors. It is also contemplated that fuel cells may be used. However, batteries and/or supercapacitors may be preferred to fuel cells because fuel cells, as well not being re-chargeable during flight, generally require or benefit from a level of ambient pressure not available at high altitude. If, however, the aircraft is deployed for lower-altitude flight or the fuel cell can be provided in a suitably pressurized environment, fuel cells could be used.

Each of the one or more propulsive gas turbine engines may comprise combustion equipment and an electrically-powered fuel pump for delivering fuel to the combustion equipment. The electrically-powered fuel pump receives its electrical power through the electrical system, for example from one or more dc distribution busses of the electrical system. Thus, the fuel pump is powered as long as there is electrical power available in the electrical system. This is in contrast to fuel pumps of conventional gas turbine engines in which the fuel pump(s) are powered as long as the spools of the gas turbine are rotating. For example, a fuel pump is conventionally either mechanically driven by a spool of the engine (e.g. through an accessory gearbox (AGB)) or powered by the output a dedicated variator arrangement which has mechanical power from a spool of the engine as one if its inputs (e.g. through an AGB). Utilizing an electrically-powered fuel pump, however, allows for a reduction in weight and mechanical losses due to a reduction in mechanical parts, and may allow faster and more precise control over the delivery of fuel to the combustion equipment.

The control system may be configured to control the one or more gas turbine engines and the electrical system to maintain a state of charge of the electrical energy storage system above a pre-defined level sufficient to power the electrically-powered fuel pump during the time period ΔT.

In some embodiments, the first spool is a low pressure (LP) spool and the second spool is a high pressure (HP) spool. The propulsive gas turbine engine may further comprise a fan mechanically coupled with and driven by the first (LP) spool, possibly via reduction gearbox whereby the fan rotates at a lower speed than the first (LP) spool. In other examples the engine is a three-spool engine with a HP spool, an intermediate pressure (IP) spool and an LP spool driving a fan. In this case the second spool may be the HP spool and the first spool may the IP spool.

A ratio, P<NUM>/T, equal to the maximum electrical power generation rating of the first electric machine divided by the maximum thrust rating of the gas turbine engine, may be greater than or equal to <NUM> WN-<NUM>. The thrust rating is a dry thrust rating (i.e. without the use of reheat/afterburner, if such a system is present in the engine). The first spool may be an intermediate pressure (IP) spool of the gas turbine engine. The term "intermediate pressure spool" will be understood to refer to a spool which is neither the highest pressure nor the lowest pressure spool of the engine. In other embodiments, the first spool is a low pressure (LP) spool of the gas turbine engine. The term "low pressure spool" will be understood to refer to the lowest pressure spool of the gas turbine engine. In some embodiments, the LP spool drives a fan via a reduction gearbox so that the fan rotates at a lower speed than the LP spool. The value of the ratio, P<NUM>/T, may be significantly higher than in conventional gas turbine engines. The value of the ratio, P<NUM>/T, may be greater than <NUM> WN-<NUM>, greater than <NUM> WN-<NUM>, or even greater than <NUM> WN-<NUM>. The value of the ratio, P<NUM>/T, may be less than <NUM> WN-<NUM>, less than <NUM> WN-<NUM>, less than <NUM> WN-<NUM>, or less than <NUM> WN-<NUM>. The value of the ratio, P<NUM>/T, may be between <NUM> WN-<NUM> and <NUM> WN-<NUM> or between <NUM> WN-<NUM> and <NUM> WN-<NUM>. In one embodiment the value of the ratio, P<NUM>/T, is between <NUM> WN-<NUM> and <NUM> WN-<NUM>.

A ratio, P<NUM>/T, equal to the maximum electrical power generation rating of the second electric machine divided by the maximum thrust rating of the gas turbine engine, may be greater than or equal to <NUM> WN-<NUM>. The thrust rating is a dry thrust rating (i.e. without the use of reheat/afterburner, if such a system is present in the engine). The second spool may be a high pressure (HP) spool of the gas turbine engine. The term "high pressure spool" will be understood to refer to the highest pressure spool of the engine. The value of the ratio, P<NUM>/T, may be significantly higher than in conventional gas turbine engines. The value of the ratio, P<NUM>/T, may be greater than <NUM> WN-<NUM>, greater than <NUM> WN-<NUM>, greater than <NUM> WN-<NUM> or even greater than <NUM> WN-<NUM>. The value of the ratio, P<NUM>/T, may be less than <NUM> WN-<NUM>, less than <NUM> WN-<NUM>, less than <NUM> WN-<NUM>, or less than <NUM> WN-<NUM>. The value of the ratio, P<NUM>/T, may be between <NUM> WN-<NUM> and <NUM> WN-<NUM>, between <NUM> WN-<NUM> and <NUM> WN-<NUM>, between <NUM> WN-<NUM> and <NUM> WN-<NUM> or between <NUM> WN-<NUM> and <NUM> WN-<NUM>. In one embodiment the value of the ratio, P<NUM>/T, is between <NUM> WN-<NUM> and <NUM> WN-<NUM>. In another embodiment the value of the ratio, P<NUM>/T, is between <NUM> WN-<NUM> and <NUM> WN-<NUM>. In another embodiment the electric machines of the first and second spool are equally rated, e.g. P1/T and P2/T may have equal or substantially equal values greater than or equal to <NUM> WN-<NUM>.

A ratio defined as a combined rated power of all of the one or more electrical machines mechanically coupled with the spools of the gas turbine engine divided by the maximum rated thrust of the gas turbine engine may be greater than or equal to <NUM> W/N. The thrust rating is a dry thrust rating (i.e. without the use of reheat/afterburner, if such a system is present in the engine). This may be considerably higher than the equivalent ratio for conventional gas turbine engines. The increased ratio may be achieved by providing electrical machines coupled with both the first spool and the second spool (i.e. dual spool power generation). The higher ratio may, for example, allow rapid charging of the energy storage system and an improved restart envelope. The ratio may be greater than <NUM> W/N, for example greater than or equal to <NUM> W/N, greater than or equal to <NUM> W/N, greater than or equal to <NUM> W/N, greater than or equal to <NUM> W/N, or even greater than or equal <NUM> W. The ratio may be less than <NUM> W/N. In one group of embodiments, the ratio is between <NUM> W/N and <NUM> W/N, and is preferably between <NUM> W/N and <NUM> W/N. In one embodiment, the ratio is between <NUM> W/N and <NUM> W/N.

The term "sub-machine" of an electric machine will be understood by those skilled in the art to refer to one of a plurality of functionally separate electric machines which share some common structure. In other words, each sub-machine is capable of generating separate electrical power or torque (depending on the mode of the operation, i.e. generator mode or motor mode), but each one of the plurality of sub-machines is not totally physically independent from the other sub-machines and thus all of the sub-machines must be deployed together. For example, two sub-machines of an electric machine may share a common rotor but have completely independent stators; they may some common stator structure (e.g. a common back iron or yoke) but have separate stator field windings; or they may be integrated together in a common casing with common support structures and cooling systems. Thus, an electric machine with two sub-machines may be considered to be an integrated arrangement of two functionally separate electric machines.

Each propulsive gas turbine engine may be of any suitable type, for example of the ducted fan type (e.g. a turbofan) or the unducted, open rotor, type.

The control system may take the form of a controller or one or more controllers, each of which may be implemented in hardware, software of a combination of the two. The control system may comprise one or more functional modules of a wider control system. In some embodiments, the control system is or is part of an Electronic Engine Controller (EEC), Engine Control Unit (ECU) or Full Authority Digital Engine Controller (FADEC).

It will be understood that the low-pressure compressor <NUM>, the low-pressure turbine <NUM> and the interconnecting shaft <NUM> together form the low-pressure spool. Similarly, the high-pressure compressor <NUM>, the high-pressure turbine <NUM> and the interconnecting shaft <NUM> together form the high-pressure spool.

There are four planet gears <NUM> illustrated, although it will be apparent to the skilled reader that more or fewer planet gears <NUM> may be provided.

It will be appreciated that the arrangement shown in <FIG> and <FIG> is by way of example only. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used.

Accordingly, although not within the scope of the claims, the examples extend to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in <FIG> has a split flow nozzle <NUM>, <NUM> meaning that the flow through the bypass duct <NUM> has its own nozzle <NUM> that is separate to and radially outside the core engine nozzle <NUM>. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct <NUM> and the flow through the core <NUM> are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, other examples my apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine <NUM> may not comprise a gearbox <NUM>.

<FIG> is a schematic illustration of an aircraft <NUM> which includes a power and propulsion system <NUM>.

The power and propulsion system <NUM> includes a propulsive gas turbine engine <NUM>. The engine <NUM> has two spools: a first spool, which in this case is a low-pressure spool having a low-pressure compressor <NUM>; a low-pressure turbine <NUM>; and an interconnecting shaft <NUM>, and a second spool, which in this case is a high-pressure spool having high-pressure compressor <NUM>; a high-pressure turbine <NUM>; and an interconnecting shaft <NUM>. The engine could alternatively be of the three-spool type and/or could include a gearbox as described above with reference to <FIG>. The engine <NUM> may be of the low- or high-bypass turbofan type or another type, for example the open rotor type.

The propulsive gas turbine engine <NUM> includes two electric machines <NUM>, <NUM>: a first electric machine <NUM> mechanically coupled with its first spool and a second electric machine <NUM> coupled with its second spool. Each of the machines <NUM>, <NUM> is operable as both a motor and as a generator. Thus, each machine <NUM>, <NUM> can either drive rotation of its spool to increase its speed, or extract mechanical power from its spool and generate electrical power therefrom.

In a preferred embodiment one or both of the electric machines <NUM>, <NUM> are coaxially coupled with the shafts of the spools and embedded within the core <NUM> of the engine <NUM>. The electric machines <NUM>, <NUM> could, however, be coupled with the spools in another manner, for example through an auxiliary gearbox using a radial/tower shaft arrangement.

In this embodiment each electric machine <NUM>, <NUM> comprises two sub-machines: the first electric machine <NUM> comprises a first sub-machine <NUM>' and a second sub-machine <NUM>"; and the second electric machine <NUM> comprises a first sub-machine <NUM>' and a second sub-machine <NUM>". The term "sub-machine" will be understood to refer to one of a plurality of functionally separate electric machines which are physically integrated together to some extent. For example, the electric machine <NUM> may be a dual-wound machine having a single, common rotor and a single stator structure featuring two independent sets of field windings, with the two independent sets of windings forming the two sub-machines <NUM>', <NUM>". Alternatively, each of the two sub-machines <NUM>', <NUM>" may have its own stator and its own rotor axially spaced part from the rotor and stator of the other sub-machine, with some common structure such as a common casing or mounting arrangement. Other sub-machine arrangements will occur to those skilled in the art. The overall power rating of each electric machine is preferably equally or approximately equally split between its two sub-machines, though this need not necessarily be the case: a split of <NUM>-<NUM>% and <NUM>-<NUM>% could be used, for example.

Each electric machine <NUM>, <NUM> can be of any suitable type known in the art. In a preferred embodiment each machine <NUM>, <NUM> is of the permanent magnet type, but induction or switched-reluctance machines could also be used, for example. The machines <NUM>, <NUM> are preferably of the radial flux or transverse flux type, but other arrangements could be used.

The power and propulsion system <NUM> further includes an electrical system <NUM>. The electrical system <NUM> includes one or more electrical distribution busses <NUM>, <NUM> which are electrically connected with the electric machines <NUM>, <NUM> of the engine <NUM>. In this way, electrical power can be received from, delivered to and transferred between the electric machines <NUM>, <NUM>.

In the illustrated embodiment the electrical system <NUM> is a dc electrical system with one or more dc electrical distribution busses <NUM>, <NUM> which interface with the electric machines <NUM>, <NUM> through ac-dc power electronics converters <NUM>, <NUM>. The use of a dc electrical system allows a single bus to simultaneously receive electrical power from electric machines driven to rotate at different speeds, allowing for what may be referred to as 'dual spool power generation'. This means that significant amounts of electrical power can be generated from one engine <NUM> even at low-power engine operating points. Additionally or alternatively, the ability to simultaneously generate power from two spools of the same engine may reduce the impact on the engine operating point. This allows the engine surge margin to be maintained whilst a relatively large amount of electrical power is generated to, for example, simultaneously charge an energy storage system <NUM> and power engine and platform electrical loads <NUM>, <NUM>, <NUM>. In other embodiments an ac electrical system may be used.

Each electric machine <NUM>, <NUM> is connected with the dc electrical system <NUM> via a set of one or more ac-dc converters. Specifically, the first electric machine <NUM> has a first set of bidirectional ac-dc converters <NUM>, including a first converter <NUM>' for its first sub-machine <NUM>' and a second converter <NUM>" for its second sub-machine <NUM>". Likewise, the second electric machine <NUM> has a second set of bidirectional ac-dc converters <NUM>, including a third converter <NUM>' for its first sub-machine <NUM>' and a fourth converter <NUM>" for its second sub-machine <NUM>". Any suitable ac-dc converter topology may be used, for example H-bridges accompanied by appropriate filters.

In the illustrated embodiment, the dc-sides of the ac-dc converters <NUM>, <NUM> are connected with the dc busses <NUM>, <NUM> so as to provide re-configurability and fault tolerance, in addition to simultaneous dual-spool power generation. Specifically, since each electric machine <NUM>, <NUM> comprises two sub-machines, each electric machine can be connected with each of the two busses <NUM>, <NUM>. The first sub-machine <NUM>' of the first electric machine <NUM> is connectable with the first dc bus <NUM>, the second sub-machine <NUM>" of the first electric machine <NUM> can be connected with the second dc bus <NUM>, the first sub-machine <NUM>' of the second electric machine <NUM> can be connected with the first dc bus <NUM>, and the second sub-machine <NUM>" of the second electric machine <NUM> can be connected with the second dc bus <NUM>.

Through the electrical system <NUM>, the power and propulsion system <NUM> provides electrical power to various electrical loads distributed about the engine <NUM> and the aircraft platform <NUM>. Platform electrical loads <NUM> may include, for example, one or more of: lighting, cabin environmental control systems such as heating systems, wing anti-icing systems, various actuators and the like. The engine electrical loads <NUM>, <NUM> may include an electrically-powered fuel pump <NUM>. Other engine electrical loads <NUM> may include, for example, an electric nacelle anti-icing system, an electrically-powered oil pump, or an electric cabin blower system.

Some of the electrical loads may be considered critical electrical loads. For example, if the engine <NUM> includes an electrically powered fuel pump <NUM> which delivers fuel to the combustion equipment <NUM> of the engine <NUM>, this will be considered a critical load since a failure of or loss of power to the fuel pump will result in the engine <NUM> shutting down.

The power and propulsion system <NUM> further includes an energy storage system (ESS) <NUM>. In preferred embodiments the ESS <NUM> takes the form of a rechargeable battery pack or module <NUM>, formed from lithium-ion cells or cells of another suitable type. As illustrated by the dashed lines, the ESS is connected with the electrical system <NUM>, possibly via a dc-dc power electronics converter <NUM> for conditioning the power and matching it to voltage of dc electrical system. In this embodiment it can be seen that the ESS <NUM> is connected with each of the dc distribution busses <NUM>, <NUM>.

The ESS <NUM> may be used as an electrical power source or sink for a variety of different purposes. For example, the ESS <NUM> may provide power to, or sink power from, the spool-coupled electric machines <NUM>, <NUM> in order to manage the surge margin of engine, especially during engine transients such as accelerations and decelerations. The ESS <NUM> may also be used to power one or more of the electric machines <NUM>, <NUM> as part of an electric start procedure. In accordance with the present invention, the ESS <NUM> is used to provide a time period, ΔT, during which the power demands of the electrical loads <NUM>, <NUM>, <NUM> can be reduced and the operating point of the engine <NUM> modified in the event of a fault or failure which results in a drop in the electrical power generation capability of the system <NUM>.

The aircraft <NUM> may further include an Auxiliary Power Unit (APU) <NUM>. In some embodiments, one or more electric machines (not shown) driven by the APU <NUM> are used as an electrical power source to replace or supplement the power provided by the ESS <NUM> during an electric start of the engine <NUM>. In other embodiments, however, the APU <NUM> may not be used for this purpose, or may be entirely omitted from the aircraft <NUM> to reduce weight. In some instances, for example where the ESS <NUM> is insufficiently charged and the APU is omitted or unable to provide power, a ground cart may be used to provide electrical power to start the engine <NUM>.

An APU <NUM> is generally not capable of being started during flight. Thus, the aircraft may also further include a Ram Air Turbine (RAT) <NUM> or other emergency source of electrical power. In case of a loss of electrical power, the RAT <NUM> may be deployed, for example to charge the ESS <NUM> to allow an inflight restart attempt. In preferred embodiments, however, the RAT <NUM> may be omitted entirely to reduce aircraft weight, with the ESS <NUM> providing sufficient engine restart capability. In some embodiments a portion of the ESS <NUM> may be dedicated to providing electrical power during restart attempts so that the capability is always available. Additionally or alternatively a state of charge of the ESS <NUM> may be maintained above a threshold level so that the capability is always available.

The power and propulsion system <NUM> further includes a control system <NUM>. The control system <NUM>, which can take any suitable form including a FADEC and may include one or more controllers and/or one or more functional modules, provides control of the engine <NUM>, including the fuel pump <NUM> and the electrical machines <NUM>, <NUM>; and the electrical system <NUM>, including the ESS <NUM> and the power electronics <NUM>, <NUM>, <NUM>. The control system <NUM> may, amongst other things, control: the configuration of the electrical system <NUM>; the modes in which the electric machines <NUM>, <NUM> of the engine <NUM> operate; the delivery of fuel into the combustion equipment <NUM> of the engine <NUM>; the mode of operation of the ESS <NUM>, the APU <NUM> and/or the RAT <NUM>; and parameters of the power electronics (e.g. switching frequencies and duty cycles of the semiconductor switches). An inflight method of controlling the power and propulsion system <NUM>, at least partially performed under the control of the control system <NUM>, is described below with reference to <FIG>.

<FIG> is a schematic illustration of an aircraft <NUM> with a power and propulsion system <NUM>. The power and propulsion system <NUM> is substantially the same as that of <FIG>, except that it includes two propulsive gas turbine engines 110a, 110b instead of one.

In the illustrated embodiment, the dc busses 123a, 124a associated with the first engine 110a are electrically connected with the dc busses 123b, 124b associated with the second engine 110b, optionally via selectively openable and closable bus ties <NUM>, <NUM>. This allows electrical power generated by the electric machines <NUM>, <NUM> of one engine (e.g. engine 110a) to be transferred to the other engine (e.g. engine 110b). This generally increases fault tolerance in the power generation system, and allows, for example, critical loads associated with the second engine 110b (e.g. its fuel pump <NUM>) to be powered by electrical power generated by the first engine 110a in the event of a partial or total loss of electrical power generation in the second engine 110b.

Compared with existing aircraft power and propulsion systems, the power and propulsion systems <NUM> of the present disclosure (including both the single engine embodiment of <FIG> and the multi-engine embodiment of <FIG>) may have electric machines <NUM>, <NUM> sized and designed so that a much greater amount of electrical power can be generated from the engine spools. Combined with dual spool power generation, this means various functions such as the simultaneous management of engine surge margin and charging of the ESS <NUM> are possible over a wide range of engine operating points.

Table <NUM> illustrates exemplary peak powers for the first machine <NUM> (coupled with the first spool, which may be a low-pressure or intermediate-pressure spool) and the second electric machine <NUM> (coupled with the second spool, which may be a high-pressure spool). The values are expressed as ratios, the divider being the peak engine thrust. For the avoidance of doubt, the 'peak thrust' is the peak dry thrust, which refers to the peak thrust without the use of any afterburner or reheat.

For both the electric machine power and the engine thrust, the term 'peak' will be readily understood by those skilled in the art to refer to the 'rated' values, i.e. the maximum values for which the electric machine or engine are designed to operate in without causing damage to the components. As explained above, each electric machine <NUM>, <NUM> may include multiple sub-machines <NUM>', <NUM>", <NUM>', <NUM>". In this case, the peak power an electric machine (e.g. electric machine <NUM>) is the sum of the peak powers of each of its sub-machines <NUM>', <NUM>".

Table <NUM> illustrates exemplary maximum values of the fraction of the power extracted from the spools of an engine (e.g. <NUM>, 110a) by its electric machines during flight. In other words, the maximum value of the following percentage to occur during a flight of the aircraft <NUM>: <MAT>.

It will be appreciated that whilst the above relates to a two-spool engine, the same percentage can be calculated for a three-spool engine. Furthermore, it will be appreciated that the absolute values of the spool power and electrical power generation are platform-dependent and will increase or decrease depending on the size of the platform and its engines.

Compared with existing aircraft power and propulsion systems, the ESS <NUM> of the power and propulsion system <NUM> of the present disclosure may be sized and designed so as to provide a greater amount of electrical power relative to the size of the platform <NUM>. Table <NUM> illustrates exemplary ESS energy storage capacities as well peak power and propulsion system thrusts (i.e. the sum of the peak thrusts of all of the propulsive engines of the platform). A ratio, defined as the total energy storage capacity divided by the maximum rated thrust of the power and propulsion system is also provided.

Various modifications and alternatives to the specific embodiments illustrated in <FIG> and <FIG> will occur to those skilled in the art. For example:.

<FIG> is a flow chart illustrating a method <NUM> of operating an aircraft power and propulsion system <NUM> during a flight. Steps of the method <NUM> may be performed automatically or by a pilot of the aircraft <NUM> or a combination of the two, and may be performed under the control of a control system, for example the control system <NUM> of the power and propulsion system <NUM> of <FIG> and <FIG>. In the method <NUM>, a fault or failure occurs in the power and propulsion system <NUM> (e.g. in an engine <NUM> and/or the electrical system <NUM>) and this results in sudden drop in the power generation level in the power and propulsion system.

The method <NUM> begins at <NUM> where the aircraft <NUM> takes off and climbs, for example to a preferred cruising altitude. In some embodiments, the ESS <NUM> discharges to the electrical system <NUM> during the take-off and/or climb procedure <NUM> and power is delivered from the electrical system <NUM> to one or more of the electric machines <NUM>, <NUM> of the propulsive gas turbine engine(s) <NUM> which operate as motor(s). The motor(s) <NUM>, <NUM> drive rotation of the engine spool(s) so as to increase their speeds and increase the engine thrust. This may, for example, reduce the amount of fuel that is fuel that is consumed during take-off and climb, and/or reduce the amount of noise produced by the engine.

At <NUM>, with the aircraft <NUM> and its propulsion system <NUM> operating at a desired operating point, one or more of the electric machine(s) <NUM>, <NUM> of the engine(s) operate in generator mode to extract mechanical power from the spool(s) and generate electrical power therefrom. The electrical power is delivered to the electrical system <NUM>, where it is utilized to power a plurality of electrical loads connected with the electrical system <NUM>. The electrical loads can include engine electrical loads <NUM>, <NUM> and platform electrical loads <NUM>. Some of the electrical power may be transferred from the electric machine (e.g. <NUM>) of one spool to the electric machine (e.g. <NUM>) of another spool of the engine, or a spool of a different engine in multi-engine embodiments.

In some embodiments, the electrical loads include one or more critical electrical loads, for example one or more engine electric fuel pumps <NUM>, whose operation is critical to the operation of the overall power and propulsion system <NUM>. Thus, the electrical loading of the power and propulsion system <NUM> may have a relatively high level of criticality.

The amount of power extracted from engine spools will vary during the course of the flight, but will generally be relatively high. The amount of power extracted from one engine, expressed as a fraction of the total shaft power of the engine, is higher than <NUM>% or even <NUM>% at some points during the normal operation of the system <NUM>.

At <NUM>, during the normal operation of the aircraft <NUM> and its propulsion system, the ESS <NUM> may be intermittently discharged to the electrical system <NUM>. For example, it may be desirable to quickly provide electrical power to one or more of the electric machines <NUM>, <NUM> of the engine(s) in order to operate them as motors to manage engine surge margin during engine transients (e.g. accelerations or decelerations) or aircraft manoeuvres (e.g. turns and climbs).

At optional step <NUM>, during the normal operation of the aircraft <NUM> and its propulsion system <NUM>, the state of charge of the ESS <NUM> is maintained at or above a predefined level. The specific level above which the charge is maintained will depend on the specifics of the platform and the application requirements, but in accordance with the present disclosure will be high enough to power one or more and optionally all of the electrical loads <NUM>, <NUM>, <NUM> during a time period ΔT. In some embodiments, the one or more electrical loads includes at least the critical electrical loads (e.g. one or both of an electric fuel pump <NUM> and an electric oil pump). The time period may be of the order of several seconds (e.g. <NUM> to <NUM> seconds) or possibly as high as several minutes (e.g. one or two minutes).

In some embodiments, maintaining the state of charge of the ESS <NUM> above the pre-defined level comprises recharging the ESS <NUM> following the initial discharge during take-off and climb, i.e. at step <NUM>. For example, the control system <NUM> may be configured to begin or complete a recharge of the ESS <NUM> up to at least the predefined level within a predetermined length of time after take-off.

Additionally or alternatively, maintaining the state of charge of the ESS <NUM> may involve recharging the ESS <NUM> up to at least the predefined level following an intermittent discharge such as the one described above with reference to step <NUM>. As another example, the control system <NUM> may be configured so that it does not allow the ESS <NUM> to discharge below the predefined level, for example by limiting intermittent discharges of the type described with reference to step <NUM> where a discharge would take the charge level of the ESS below the threshold.

In other embodiments, a portion of the ESS <NUM> may be dedicated to certain functions, particularly those described below with reference to steps <NUM>-<NUM> and <NUM>-<NUM>. For example, one or more battery cells or modules of the ESS <NUM> may be used only for a certain set of functions and may be never be used for more general purposes such as intermittent discharge. If the ESS <NUM> does include a portion dedicated to said functions, the maintenance of the state of charge of the ESS <NUM> may be avoided.

At step <NUM>, the control system <NUM> determines a condition to the effect that a fault or failure in one or more of the gas turbine engines <NUM> and/or in the electrical system <NUM> has occurred, and that the amount of electrical power that is being generated by the system <NUM> has dropped to a reduced level, Pfault. Possible faults and failures which could lead to a drop in the amount of power that is generated include, but are not limited to:.

At the time of the fault, the electrical system <NUM> will have been subject to an instantaneous power demand, PD. If the instantaneous power demand prior to the fault PD is greater than the new, lower level of power generation capability, Pfault, the method <NUM> may proceed to steps <NUM>-<NUM>. If instead the instantaneous power demand prior to the fault PD is lower than the new, lower level of power generation capability, Pfault, the method <NUM> may proceed to steps <NUM>-<NUM>.

At <NUM>, responsive to the determination of the condition to the effect that there has been a fault or failure in the system <NUM> and the amount of power being generated has suddenly dropped to below the power demand level PD, the ESS <NUM> discharges and provides power to the electrical system for a time ΔT. The amount of electrical power supplied by the ESS <NUM>, Pass, may be at least equal to the deficit in power, i.e. PD - Pfault. Preferably, however, the ESS <NUM> supplies an amount of power PESS > PD - Pfault. In this way, in addition to ensuring that all electrical loads <NUM>, <NUM>, <NUM> are powered during the time period ΔT, the reliance on power generated by the electric machine(s) <NUM>, <NUM> of the engine(s) can be reduced and this may limit the impact on the engine surge margin during the time period ΔT. In other embodiments, PESS < PD - Pfault. For example, the ESS <NUM> may only provide enough electrical power to supply one or more critical electrical loads (e.g. a fuel pump <NUM> and/or oil pump) so that the only electrical loads which may be starved of power in the time period ΔT are non-critical loads.

At steps <NUM>-<NUM>, the control system <NUM> uses the time period ΔT provided by the ESS <NUM> to adjust the operating point of the power and propulsion system <NUM> so that the electrical power demand, PD, is lower than the amount of electrical power that is being generated by the power and propulsion system <NUM>. In this way, after the time period ΔT, the power and propulsion system <NUM> is at a sustainable operating point. This may be achieved by implementing either one or both of steps <NUM> and <NUM>.

At step <NUM>, the control system <NUM> controls the plurality of electrical loads <NUM>, <NUM>, <NUM> to reduce the total system power demand, PD. Preferably, the power demand, PD, is reduced to a level below Pfault. This may be achieved by completely powering off some or all of the non-critical electrical loads or adjusting the operation of some or all of the loads to states in which they consume less electrical power.

At step <NUM>, the control system controls the gas turbine engine(s) <NUM>, including the electric machine(s) <NUM>, <NUM>, to adjust the operating points of the engines(s). For example, the engine(s) may be controlled to operate at higher thrust operating points so that more electrical power can be extracted by the electric machines <NUM>, <NUM> without excessive erosion of the engine surge margin.

If instead PD < Pfault, the method proceeds from step <NUM> to step <NUM>.

At step <NUM>, responsive to the determination of the condition to the effect that there has been a fault or failure in the system <NUM> and the amount of power being generated has suddenly dropped, the ESS <NUM> discharges and provides power to the electrical system for a time ΔT. Even though the amount of power being generated by the electric machines <NUM>, <NUM> may be sufficient to power all of the loads, the use of the ESS <NUM> in this way may have the effect of softening the impact of the fault on the surge margin(s) of the engines <NUM>.

At steps <NUM>-<NUM>, the control system <NUM> uses the time period ΔT provided by the ESS <NUM> to adjust the operating point of the power and propulsion system <NUM> so that, not only is the electrical power demand PD lower than the amount of electrical power that is being generated by the power and propulsion system <NUM>, but the engine(s) have sufficient surge margin to sustain the system operating point. This may be achieved by implementing either one or both of steps <NUM> and <NUM>.

At step <NUM>, the control system <NUM> controls the plurality of electrical loads <NUM>, <NUM>, <NUM> to reduce the total system power demand, PD. This may be achieved by completely powering off some or all of the non-critical electrical loads or adjusting the operation of some or all of the loads to states in which they consume less electrical power.

In some embodiments, performing either one or both of steps <NUM>-<NUM>, or performing one or both of steps <NUM>-<NUM>, may comprise implementing a pre-defined 'limp mode'. The term `limp mode' or 'limp home mode' will be understood to refer to a pre-defined set of operating parameters for the power and propulsion system <NUM> in which critical loads and systems operate but the overall power demand is reduced. The use of a pre-defined limp mode may reduce the amount of time required by the control system <NUM> to move the power and propulsion system <NUM> to a sustainable operating point. In some embodiments, a number of different limp modes may be defined, each corresponding to one or more fault or failure modes. Thus, responsive to a diagnosis of the fault, the control system <NUM> may select and implement a suitable one of the limp modes during the time period ΔT.

Therefore, a method <NUM> is provided in which an ESS <NUM> of a power and propulsion system <NUM> is used to provide a time period ΔT following a fault during which an operating point of the power and propulsion system <NUM> can be adjusted. Such an approach may particularly effective in a platform in which there is a relatively high degree of critically in the electrical loading, as it may help mitigate the risk that critical loads lose power following the fault. Even a brief loss of power to critical loads could result in an escalation of a fault. Additionally or alternatively, the approach may be particularly effective in a platform in which the amount of mechanical power which is extracted from the engine spools to generate electrical power is relatively large, greater than <NUM>% or as high as for example <NUM>%, <NUM>% or even <NUM>% of the total spool power. Such a system <NUM> may be sensitive to the erosion of engine surge margin following a fault, and engine surge following a fault could result in an escalation of a fault.

Claim 1:
A method (<NUM>) of operating an aircraft power and propulsion system (<NUM>), the power and propulsion system (<NUM>) comprising one or more propulsive gas turbine engines (<NUM>; 110a, 110b), each engine comprising a plurality of spools (<NUM>, <NUM>, <NUM>; <NUM>, <NUM>, <NUM>) including a first spool and a second spool, a first electric machine (<NUM>) mechanically coupled with the first spool and a second electric machine (<NUM>) mechanically coupled with the second spool, wherein each of the first and second electric machines (<NUM>, <NUM>) of each of the one or more gas turbine engines comprises a first sub-machine (<NUM>', <NUM>') and a second sub-machine (<NUM>", <NUM>"); and an electrical system (<NUM>) connected with the electric machines and comprising an electrical energy storage system (<NUM>), the method (<NUM>) comprising:
operating (<NUM>) one or more of the electric machines (<NUM>, <NUM>) of the one or more gas turbine engines (<NUM>; 110a, 110b) as generators to extract, for each of the one or more gas turbine engines, greater than <NUM>% of a combined mechanical power of the plurality of the spools (<NUM>, <NUM>, <NUM>; <NUM>, <NUM>, <NUM>) and to generate electrical power therefrom;
meeting an electrical power demand, PD, of a plurality of electrical loads (<NUM>, <NUM>, <NUM>) connected with the electrical system (<NUM>) by supplying the plurality of electrical loads with electrical power generated by the one or more electric machines (<NUM>, <NUM>);
determining (<NUM>) a condition to the effect that there is a fault in one or more of the electric machines and that an amount of electrical power being generated by the power and propulsion system (<NUM>) has reduced to a lower level, Pfault; and
responsive to the determination (<NUM>):
during a time period ΔT, meeting (<NUM>, <NUM>) at least part of the electrical power demand of the plurality of electrical loads (<NUM>, <NUM>, <NUM>) by discharging the electrical energy storage system (<NUM>); and
during the time period ΔT, controlling (<NUM>) the plurality of electrical loads (<NUM>, <NUM>, <NUM>) to reduce the electrical power demand.