Patent Description:
Traditional helicopter architectures using two gas turbine engines are designed to meet operation and safety requirements. Each gas turbine engine is typically sized to provide additional (but temporary) power above rated maximum takeoff power (MTOP) in the event of failure of one of the gas turbine engines, for example.

Such conventional methods and systems have generally been considered satisfactory for their intended purpose. However, there is still a need in the art for improved power systems for rotorcraft, such as with respect to performance.

<CIT> discloses a hybrid electric system according to the preamble of claim <NUM>, <CIT> discloses an aircraft degraded operation ceiling increase using electric power boost, <CIT> discloses a system and method of augmenting power in a rotorcraft, <CIT> discloses a hybrid contingency power drive system, and <CIT> discloses a system and method for operating a multi-engine rotorcraft.

According to a first aspect, there is provided a hybrid electric system for a rotorcraft as set forth in claim <NUM>.

There is also provided a rotorcraft as set forth in claim <NUM>.

According to a further aspect, there is provide a non-transitory computer readable medium as set forth in claim <NUM>.

Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, an illustrative view of an embodiment of a system in accordance with the disclosure is shown in <FIG> and is designated by reference character <NUM>. Other embodiments and/or aspects of this disclosure are shown in <FIG>. Certain embodiments described herein can be used to provide lighter, smaller, safer, and more energy efficient power systems and rotorcraft.

Referring to <FIG>, according to the invention, a hybrid electric system <NUM> for a rotorcraft (e.g., a helicopter) includes a first thermal engine 103a (e.g., a turbine engine, an internal combustion engine), a second thermal engine 103b (e.g., a turbine engine, an internal combustion engine), and an electrical machine <NUM> (e.g., useable as a motor and a generator). The first thermal engine 103a is sized to produce a maximum first thermal engine power that is below a one-or-more-engine-inoperative (OEI) requirement power (e.g., as defined by FAA safety regulations) and the second thermal engine 103b is sized to produce a maximum second thermal engine power that is below the OEI requirement power. The one-or-more-engine-inoperative power can be a one-engine-inoperative power. Any suitable power requirement for any suitable number of engines inoperative is contemplated herein. The electrical machine <NUM> is sized to provide at least a remaining power needed to reach the OEI requirement power in an OEI state, for example.

Any suitable mechanical arrangement (e.g., one or more gear boxes) to deliver power from the power sources is contemplated herein. In certain embodiments, as shown in <FIG> the system <NUM> can include a main gear box <NUM> operatively connected to each of the first thermal engine 103a, the second thermal engine 103b, and the electrical machine <NUM>. The main gear box <NUM> can be configured to receive power from each of the power sources and combine the output to one or more rotors <NUM>, <NUM> (e.g., a main rotor <NUM> and a tail rotor <NUM>). Any suitable number of output rotors can be connected to one or more gear boxes <NUM> (e.g., a main rotor <NUM>, a tail rotor <NUM>, auxiliary rotor).

For example, as shown in <FIG>, a system <NUM>, e.g. for a multirotor aircraft, can include a first gearbox 101a for driving a first rotor 106a and a second gear box 101b for driving a second rotor 106b. As shown the first thermal engine 103a can be connected to the first gearbox 101a, the second thermal engine 103b can be connected to the second gearbox 101b, and the electrical machine <NUM> can be connected to both gearboxes 101a, 101b. Any suitable other arrangements and/or connections are contemplated herein.

The OEI requirement power can be defined by a required torque ratio (e.g., a percentage of normal takeoff torque) as appreciated by those having ordinary skill in the art, or defined as a function of a performance metric (e.g., the ability to climb, e.g., at <NUM> meters per minute (<NUM> feet per minute), for a certain period of time, e.g., <NUM> seconds or two and a half minutes) as appreciated by those having ordinary skill in the art. One having ordinary skill in the art appreciates how to determine an appropriate OEI power requirement based on the aircraft and/or applicable regulations.

Each of the thermal engines 103a, 103b can be connected to the main gear box <NUM> via a mechanical disconnect 107a, 107b. The electric motor <NUM> can also be connected to the main gear box <NUM> via a mechanical disconnect 107c. Any suitable connection type (e.g., the same or different for any or each) is contemplated herein.

The system <NUM> can include a battery <NUM> (e.g., having one or more cells) sized to allow for the OEI power requirement to be reached for at least a required OEI power time. For example, the OEI power time can be between about <NUM> seconds and about <NUM> minutes. Any other suitable time is contemplated herein (e.g., enough time to meet safety regulations). In certain embodiments, the electrical machine <NUM> can be configured to be driven by the first and/or second thermal engine 103a, 103b in an excess thermal power setting (e.g., during level cruise or any other time when desired or required) to generate electrical energy to charge the battery (e.g., to charge the battery, such as for example during cruise after transient use, for example to keep the battery at full charge or in a sufficient OEI requirement energy range). Any suitable charging scheme (e.g., to provide charge as soon as possible after use of the battery when excess power over what is demanded is available) is contemplated herein.

According to the invention, the system <NUM> includes one or more control modules <NUM> operatively connected to the first thermal engine 103a, the second thermal engine 103b, and the electrical machine <NUM> to control power output therefrom (e.g., from each source independently of the others, or combined output from two sources, or total output from all). In certain embodiments, the one or more control modules <NUM> can include a separate controller for each thermal engine and/or the electric motor (e.g., three separate controllers or any other suitable number). In certain embodiments, the one or more control modules <NUM> can include a single controller. The one or more control modules <NUM> can include any suitable computer hardware and/or software module(s) configured to perform any suitable function(s) disclosed herein (e.g., controlling torque from each engine and the electrical machine, controlling energy flow, e.g., discharging or charging the battery, controlling performance based on selected mode or flight state, etc.). Any suitable delineation of control modules (e.g., hardware and/or software is contemplated herein). The one or more control modules <NUM> can be configured to allow selection of a normal mode or a performance mode, for example, by the pilot. The one or more control modules <NUM> can be configured to determine an emergency state based on a condition of one or more of the thermal engines 103a, 103b and/or the electrical machine <NUM>.

According to the invention, in a normal take-off mode, the one or more control modules <NUM> are configured to cause output power only from the first thermal engine 103a and the second thermal engine 103b (e.g., so as to not utilize the electrical machine or energy stored in the battery). In a performance take-off mode, the one or more control modules <NUM> are configured to additionally cause output power to the main gear box <NUM> from the electrical machine <NUM> (e.g., to increase takeoff climb rate at the expense of battery charge state). In certain embodiments, the one or more control modules <NUM> can be configured to allow a discharge of the battery to power the electrical machine (for non-emergency performance enhancement or transient response) only to a OEI threshold where there is still enough energy to comply with the OEI power requirements.

In certain embodiments, in a normal cruise mode, the one or more control modules <NUM> can be configured to reduce power of one of the first thermal engine 103a and the second thermal engine 103b to idle or off (for fuel efficiency), and to operate the other of the first and second thermal engine 103a, 103b at a set thermal cruise power suitable for level flight, for example (or to any other type of flight that can be performed by a single thermal engine). The one or more control modules <NUM> can be configured such that, during cruise, if a demanded power changes to be above the set thermal cruise power (producible by the single thermal engine), the one or more control modules <NUM> can cause the electrical machine <NUM> to output an electrical machine power that is a difference of power between a current thermal power and the demanded power until a total thermal power produced by either or both of the first thermal engine 103a and/or the second thermal engine 103b reaches the demanded power. In certain embodiments, the electrical machine power ramps down as the total thermal power (e.g., the total power of both thermal engine 103a, 103b ramps up to maintain the demanded power (e.g., essentially compensating for a loss of torque/power in certain embodiments).

In certain embodiments, the one or more control modules <NUM> can be configured such that, in cruise, if the thermal engine (e.g., first thermal engine 103a) that is producing the set thermal cruise power fails, the one or more control modules <NUM> is configured to ramp up power from the idle or off thermal engine (e.g., second thermal engine 103b) to the demanded power and to produce the remaining power required using the electrical machine <NUM> (e.g., the difference of demanded power and thermal power). The one or more control modules <NUM> can be configured to have any suitable control logic (e.g., otherwise disclosed herein or appreciated by those having ordinary skill in the art in view of this disclosure), e.g., for any situation (e.g., emergency, normal, performance, or transient response use of the electric motor).

In certain embodiments, the one or more control modules <NUM> can have a control function that is dedicated to each lane. The one or more control modules <NUM> can receive signals from one or more rotorcraft systems with rotorcraft-level commands (throttle, CLP, mode select, etc.), and the one or more control modules <NUM> can calculate and transmit the power and speed commands to each individual power lane.

In accordance with at least one aspect of this disclosure, a rotorcraft (not shown) can include a rotor (not shown) and any suitable embodiment of a system as disclosed herein, e.g., system <NUM> as described above, connected to the rotor to power the rotor. Any other suitable components are contemplated herein.

A non-transitory computer readable medium comprising computer executable instructions is configured to cause a computer to perform a method for controlling power in a multiengine rotorcraft. The method includes reducing power of one of a first or second thermal engine to idle or off, and operating the other of the first or second thermal engine at a set thermal cruise power suitable for level flight. During cruise, if a demanded power changes to be above the set thermal cruise power, the method includes causing an electrical machine to output an electrical machine power that is a difference of power between a current thermal power and the demanded power until a total thermal power produced by either or both of the first thermal engine and/or the second thermal engine reaches the demanded power. The method further includes outputting power only from the first thermal engine and the second thermal engine in a normal take-off mode and additionally outputting power from the electrical machine in a performance take-off mode. In certain embodiments, the method can include ramping electrical machine power down as the total thermal power ramps up to maintain the demanded power.

Embodiments can include a hybrid electric propulsion system (HEPS) for a rotorcraft which can be configured to use an electric motor to provide OEI power in the event of engine failure. Instead of having to enter a power reserve regime that is over and above rated maximum takeoff power for a traditional engine (which can only be done for a very limited period of time without destroying the engine), the thermal engine can be configured to operate at a maximum continuous rating without the need for such a power reserve, and the electric motor can be used to backfill any required power, e.g., to provide suitable OEI power or performance enhancement and/or transient response.

In certain embodiments, the electrical machine is not necessary for normal use, but can be used for efficiency (allowing one engine to idle or completely shut off while providing transient response), emergency, and performance enhancements (e.g., mode to be selected by the pilot). In certain embodiments, the thermal engines can be smaller and the battery need not be particularly large, making the whole system lighter (including less robust mechanical connections to the gearbox as the amount of maximum torque from any single powerplant is reduced).

In certain embodiments, electrical machine can be used as an auxiliary power unit (APU), e.g., in an APU mode, to drive hydraulics, and electrical energy can be taken directly from the battery if necessary for other APU functions. Embodiments can also have the electrical machine to be placed in its own fire zone so high voltage wires can be segregated from fuel components.

In certain embodiments, in takeoff, just the thermal engines can be used unless the electric motor is intentionally selected by the pilot to provide more power, for example. In certain embodiments, during flight, the electric motor may be used in transients only (e.g., while waiting for thermal engines to ramp up), e.g., to climb over obstacles.

Certain embodiments allow a single thermal engine to be used, and to put other engine(s) at idle or off for efficient cruise. The electric motor can be used for OEI power only if needed, for example. If the operating engine fails, for example then the e-motor can be used to provide maximum power until the idle/off motor powers up. After this, the electric motor can be used as needed from then on if OEI power is needed, for example (e.g., to climb). In certain embodiments, the electric motor may only be used to make up for power only when absolutely needed (e.g., only for engine failure scenarios) which can minimize the battery size and weight.

In order to allow Category A operations for traditional two gas turbine engine-powered aircraft, each gas turbine must be rated to OEI power ratings (e.g. <NUM> second, <NUM> minute, continuous OEI ratings). Category A is a standard of design and operation outlined in helicopter certification requirements - FAR/CS29/AWM529. Category A certification is required for higher risk operations such as extended flight over remote areas without landing areas (e.g. forests, water, or urban areas). The OEI power is an auxiliary power rating above the maximum takeoff power (MTOP) that is only used in emergency and/or failure events. The helicopter maximum takeoff weight, performance, and operations are architected around the availability of OEI power in the event that one engine shuts down in flight. The safety criticality of OEI power dictates invasive and expensive checks to guarantee the power availability. Checks may include engine power assurance checks (EPAC), maintenance inspections, cycle counting, etc..

In certain embodiments, uncovering of potential latency of failures causing the electric motor to be unable to reach full power can be automated. In certain embodiments, during the flight, both thermal engines can reduce power and allow the electric motor to backfill the required torque to reach the commanded torque until the electric motor reaches the rated torque for OEI power. Lowering the power of both thermal engines allows functional capability in case a subsequent failure of a thermal engine occurs during the electric motor torque availability check (e.g., for category A type flights).

With the use of a parallel hybrid-electric propulsion system (HEPS) with three torque sources, e.g., two thermal and one electrical as shown in the embodiment of <FIG>, the propulsion system and rotorcraft can be optimized to provide improved performance without sacrificing safety. As disclosed above, a HEPS architecture can include two thermal engine-provided power sources and one electrical-provided power source (e.g. electrical motor) connected to the helicopter main gearbox (MGB), each with a mechanical disconnect. If <NUM>% maximum takeoff power (MTOP) is required for a helicopter takeoff rating, then each thermal engine can be rated to <NUM>% helicopter MTOP.

As an example, if <NUM>% power is required for the highest OEI rating, the electrical motor can be rated to the difference between the OEI rating and single engine rating (i.e. <NUM>% - <NUM>% = <NUM>% MTOP). Any other suitable numbers are contemplated herein.

The HEPS can run the combination of power sources as optimized for performance, fuel efficiency, power availability required for the phase of flight. The electric motor can provide power when required for safety, power availability, and/or performance. During takeoff, both thermal engines can provide all of the takeoff power required. When cruising, one thermal engine can provide the required cruise power while the other can run in low idle or shutdown mode to save fuel, referred to as Single Engine Operation (SEO). If the pilot requires additional power in SEO, the electrical machine can provide near instantaneous power to fill the required power availability or until the other thermal engine can ramp up to meet the power need. This provides greater power availability and minimal impact on helicopter performance with the SEO fuel savings. Rotor droop can be a performance concern for helicopter thermal engines, especially when using a free turbine configuration. The electrical machine can provide almost instantaneous power to mitigate any rotor droop and preserve helicopter performance. In cruise, either of the thermal engines could run at higher power to charge the battery via the electrical machine (operating as a generator).

In the event of a thermal engine failure, the controller or pilot can disconnect the failed engine via the mechanical disconnect, and the electrical machine can provide the power required to meet the One Engine Inoperative (OEI) power rating. In certain embodiments, the maximum continuous OEI rating can be about the same power level as MTOP. The electrical machine can be sized to provide enough power for the <NUM> second or <NUM> minute OEI power. In certain embodiments, the battery can be sized for the emergency power ratings plus additional energy for performance enhancements and emergency reserve, for example. Any suitable size is contemplated herein.

Embodiments allow each thermal engine to be rated to the takeoff power instead of the OEI power resulting in reduced weight, fewer required maintenance checks, and no thermal engine EPACs. The electrical machine can be sized to provide emergency or temporary power when required. This allows for reduce sizing of the electrical machine, motor controller, and the battery system. Embodiments can provide redundant sources of power that can be separated and segregated to mitigate from failure propagation (e.g. each power sources is in a different fire zone). With the electrical machine in a different fire zone, the high voltage wiring system can be separated from flammable fluid sources such as fuel.

In embodiments, any and/or all power sources can be designed to provide lower power thereby reducing weight, cost, fuel burn, and space in the aircraft. The thermal engine power sources can be designed for half of the total helicopter MTOP and optimized for single or dual engine cruise. The electrical machine can be designed to provide the remaining OEI power. Having two smaller thermal engines results in weight reduction, lower fuel burn, and lower emissions, for example.

Embodiments of a system can provide greater torque response due to the faster response of the electrical machine. This can allow for improved aircraft-level performance response. The faster response also enables single engine operation without sacrificing performance, safety, or fuel burn savings.

The elimination of thermal engine OEI rating reduces the operational and maintenance burden required to ensure power availability in the case of emergency. EPACs are often run as penalty runs which can be costly. The additional maintenance burden due to the safety criticality such as temperature probe inspection and blade inspection can be reduced. The battery sizing can be optimized for emergency power and performance boosts only. The battery can be recharged off of the thermal engine during the flight. Additionally, the engine output shafts to the MGB do not need to be sized for OEI power. They can be optimized and downsized for MTOP. Accordingly, the use of a parallel hybrid-electric propulsion system (HEPS) having two or more thermal engines and one or more electrical power source can increase efficiencies compared to traditional architectures and provide performance, safety, operations, and weight benefits.

A computer readable storage medium may be, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device. More specific examples (a non-exhaustive list) of the computer readable storage medium would include the following: an electrical connection having one or more wires, a portable computer diskette, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or Flash memory), an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, or a magnetic storage device.

Such a propagated signal may take any of a variety of forms, including, but not limited to, electro-magnetic or optical.

Program code embodied on a computer readable medium may be transmitted using any appropriate medium, including but not limited to wireless, wireline, optical fiber cable or RF.

Computer program code for carrying out operations for aspects of this disclosure may be written one or more programming languages, including an object oriented programming language such as Java, Smalltalk, C++ or the like and conventional procedural programming languages, such as the "C" programming language or similar programming languages.

Aspects of this disclosure may be described above with reference to flowchart illustrations and/or block diagrams of methods, apparatus (systems) and computer program products according to embodiments of this disclosure. These computer program instructions may be provided to a processor of a general purpose computer, special purpose computer, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, create means for implementing the functions/acts specified in any flowchart and/or block diagram block or blocks.

When separating items in a list, "or" or "and/or" shall be interpreted as being inclusive, i.e., the inclusion of at least one, but also including more than one, of a number or list of elements, and, optionally, additional unlisted items.

Claim 1:
A hybrid electric system for a rotorcraft, comprising:
a first thermal engine (103a);
a second thermal engine (103b); and
an electrical machine (<NUM>), wherein the first thermal engine (103a) is sized to produce a maximum first thermal engine power that is below a one-or-more-engine-inoperative (OEI) requirement power, the second thermal engine (103b) is sized to produce a maximum second thermal engine power that is below the OEI requirement power, and the electrical machine (<NUM>) is sized to provide at least a remaining power needed to reach the OEI requirement power in an OEI state, characterised in that the hybrid electric system further comprises:
one or more control modules (<NUM>) operatively connected to the first thermal engine (103a), the second thermal engine (103b), and the electrical machine (<NUM>) to control power output therefrom, wherein:
in a normal take-off mode, the one or more control modules (<NUM>) are configured to cause output power only from the first thermal engine (103a) and the second thermal engine (103b); and
in a performance take-off mode, the one or more control modules (<NUM>) are configured to additionally cause output power from the electrical machine (<NUM>).