Patent Description:
A conventional commercial aircraft generally includes a fuselage, a pair of wings, and a propulsion system that provides thrust. The propulsion system includes at least two aircraft engines, such as turboprop or turboshaft engines. Each engine can be mounted to a respective portion of the aircraft, such as in a suspended position beneath a wing.

<CIT> relates to systems and methods of power allocation for a hybrid electric architecture. <CIT> relates to a propulsion system for an aircraft. <CIT> also relates to a propulsion system for an aircraft.

One aspect of the present disclosure relates to an electrical system for an aircraft in accordance with appended claim <NUM>.

Another aspect of the present disclosure relates to a method of powering an aircraft system in accordance with appended claim <NUM>.

These and other features, aspects and advantages of the present disclosure will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate aspects of the disclosure and, together with the description, serve to explain the principles of the disclosure.

A full and enabling disclosure of the present description, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended FIGS. , in which:.

Example aspects of the present disclosure are directed to an electrical system for an aircraft. An aircraft electrical system that can leverage the available electric power system so that it can be used for multiple aircraft systems with high power demand (e.g., flight controls, gallery loads, and electrical actuators, propulsion systems, environmental control systems, etc.) congruently can improve aircraft performance.

The disclosure describes how an engine, such as, but not limiting to, a turboprop engine or a turbo shaft engine can be utilized to generate power from the engine with multiple spools for a first load, a second load, and a third load, which can be the high-power demand aircraft systems. The optimization of power allocation can be achieved by intelligent controllers such as a Propulsion System Controller (PSC), a Bus Power Control Unit (BPCU), or a combination thereof. Both these controllers can interface with at least one Full Authority Digital Engine Control (FADEC). These controllers (PSC, BPCU, FADEC, or a combination thereof) can work together to command the power resources for each load to ensure full flight envelope performance.

As used herein, the term "set" or a "set" of elements can be any number of elements, including only one. As used herein, the terms "axial" or "axially" refer to a dimension along a longitudinal axis of an engine or along a longitudinal axis of a component disposed within the engine. The term "forward" used in conjunction with "axial" or "axially" refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term "aft" used in conjunction with "axial" or "axially" refers to a direction toward the rear or outlet of the engine relative to the engine centerline.

Additionally, as used herein, the terms "radial" or "radially" refer to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a center longitudinal axis of the engine and an outer engine circumference. The use of the terms "proximal" or "proximally," either by themselves or in conjunction with the terms "radial" or "radially," refers to moving in a direction toward the center longitudinal axis, or a component being relatively closer to the center longitudinal axis as compared to another component.

Also as used herein, while sensors can be described as "sensing" or "measuring" a respective value, sensing or measuring can include determining a value indicative of or related to the respective value, rather than directly sensing or measuring the value itself. The sensed or measured values can further be provided to additional components. For instance, the value can be provided to a controller module or processor, and the controller module or processor can perform processing on the value to determine a representative value or an electrical characteristic representative of said value.

Additionally, while terms such as "voltage", "current", and "power" can be used herein, it will be evident to one skilled in the art that these terms can be interrelated when describing aspects of the electrical circuit, or circuit operations.

All directional references (e.g., radial, axial, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise) are only used for identification purposes to aid the reader's understanding of the disclosure, and do not create limitations, particularly as to the position, orientation, or use thereof. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to each other. In non-limiting examples, connections or disconnections can be selectively configured to provide, enable, disable, or the like, an electrical connection between respective elements. Non-limiting example power distribution bus connections or disconnections can be enabled or operated by way of switching, bus tie logic, or any other connectors configured to enable or disable the energizing of electrical loads downstream of the bus. Additionally, as used herein, "electrical connection" or "electrically coupled" can include a wired or wireless connection. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary.

Additionally, as used herein, a "controller" or "controller module" can include a component configured or adapted to provide instruction, control, operation, or any form of communication for operable components to effect the operation thereof. A controller module can include any known processor, microcontroller, or logic device, including, but not limited to: field programmable gate arrays (FPGA), an application specific integrated circuit (ASIC), a full authority digital engine control (FADEC), a proportional controller (P), a proportional integral controller (PI), a proportional derivative controller (PD), a proportional integral derivative controller (PID controller), a hardware-accelerated logic controller (e.g. for encoding, decoding, transcoding, etc.), the like, or a combination thereof. Non-limiting examples of a controller module can be configured or adapted to run, operate, or otherwise execute program code to effect operational or functional outcomes, including carrying out various methods, functionality, processing tasks, calculations, comparisons, sensing or measuring of values, or the like, to enable or achieve the technical operations or operations described herein. The operation or functional outcomes can be based on one or more inputs, stored data values, sensed or measured values, true or false indications, or the like. While "program code" is described, non-limiting examples of operable or executable instruction sets can include routines, programs, objects, components, data structures, algorithms, etc., that have the technical effect of performing particular tasks or implement particular abstract data types. In another non-limiting example, a controller module can also include a data storage component accessible by the processor, including memory, whether transient, volatile or non-transient, or non-volatile memory.

Additional non-limiting examples of the memory can include Random Access Memory (RAM), Read-Only Memory (ROM), flash memory, or one or more different types of portable electronic memory, such as discs, DVDs, CD-ROMs, flash drives, universal serial bus (USB) drives, the like, or any suitable combination of these types of memory. In one example, the program code can be stored within the memory in a machine-readable format accessible by the processor. Additionally, the memory can store various data, data types, sensed or measured data values, inputs, generated or processed data, or the like, accessible by the processor in providing instruction, control, or operation to effect a functional or operable outcome, as described herein. In another non-limiting example, a control module can include comparing a first value with a second value, and operating or controlling operations of additional components based on the satisfying of that comparison. For example, when a sensed, measured, or provided value is compared with another value, including a stored or predetermined value, the satisfaction of that comparison can result in actions, functions, or operations controllable by the controller module. As used, the term "satisfies" or "satisfaction" of the comparison is used herein to mean that the first value satisfies the second value, such as being equal to or less than the second value, or being within the value range of the second value. It will be understood that such a determination may easily be altered to be satisfied by a positive/negative comparison or a true/false comparison. Example comparisons can include comparing a sensed or measured value to a threshold value or threshold value range.

As used herein, an "essential" electrical load can be a subset of one or more electrical loads of a power distribution system or architecture classified or categorized as "essential" or "critical" to the operation of the power architecture, vehicle, or another system. In one non-limiting aspect, an "essential" electrical load can be critical to flight operations of an aircraft or critical aircraft systems, and can be defined by relevant federal aircraft regulations or relevant industry standards.

Reference now will be made in detail to aspects of the present disclosure, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the disclosure, not limitation of the disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present disclosure without departing from the scope of the disclosure. For instance, features illustrated or described as part of one aspect can be used with another aspect to yield a still further embodiment. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims.

As used herein, the term "about," when used in reference to a numerical value is intended to refer to within <NUM>% of the numerical value. As used in the specification and the appended claims, the singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

<FIG> depicts a top-down schematic view of an example aircraft <NUM> according to example aspects of the present disclosure. The aircraft defines a longitudinal centerline <NUM> that extends there through, a vertical direction, a lateral direction, L, a forward end <NUM>, and an aft end <NUM>. Moreover, the aircraft <NUM> defines a mean line extending between the forward end <NUM> and aft end <NUM> of the aircraft <NUM>. As used herein, the "mean line" refers to a midpoint line extending along a length of the aircraft <NUM>, not taking into account the appendages of the aircraft <NUM> (such as the wings <NUM> and stabilizers discussed below).

Moreover, the aircraft <NUM> includes a fuselage <NUM>, extending longitudinally from the forward end <NUM> of the aircraft towards the aft end <NUM> of the aircraft <NUM>, and a pair of wings <NUM>. As used herein, the term "fuselage" generally includes all of the body of the aircraft <NUM>, such as an empennage of the aircraft <NUM>. The first such wings <NUM> extends laterally outwardly with respect to the longitudinal centerline <NUM> from a port side <NUM> of the fuselage <NUM> and the second of such wings <NUM> extends laterally outwardly with respect to the longitudinal centerline <NUM> from a starboard side <NUM> of the fuselage <NUM>. Each of the wings <NUM> for the example aspect depicted includes one or more leading edge flaps <NUM> and one or more trailing edge flaps <NUM>. The aircraft <NUM> further includes a vertical stabilizer having a rudder flap for yaw control, and a pair of horizontal stabilizers <NUM>, each having an elevator flap <NUM> for pitch control. The fuselage <NUM> additionally includes an outer surface or skin <NUM>. It should be appreciated however, that in other example aspects of the present disclosure, the aircraft <NUM> may additionally or alternatively include any other suitable configuration of stabilizer that may or may not extend directly along the vertical direction or horizontal/lateral direction L.

The example aircraft <NUM> of <FIG> includes a system <NUM>. The example system <NUM> includes one or more aircraft engines. For example, the aspect depicted includes a plurality of aircraft engines, shown as engines <NUM>, <NUM>, each configured to be mounted to the aircraft <NUM>. More specifically, in the aspects of the disclosure depicted, the aircraft engines <NUM>, <NUM> are configured as gas turbine engines <NUM>, <NUM>, or rather as turboprop or turboshaft engines, attached to and suspended beneath the wings <NUM> in an under-wing configuration.

In one non-limiting example aspect of the disclosure, the system <NUM> can further includes one or more electric generators <NUM> kinetically connected with the engines <NUM>, <NUM>. For example, one or both of the engines <NUM>, <NUM> may be configured to provide mechanical power from a rotating shaft (such as an LP shaft or HP shaft) to the electric generators <NUM>. Although depicted schematically outside the respective engines <NUM>, <NUM>, in certain aspects, the electric generators <NUM> may be positioned within or proximate to a respective engine <NUM>, <NUM>. Additionally, the electric generators <NUM> may be configured to convert the mechanical power to electrical power. For the aspect depicted, the system <NUM> includes at least one electric generator <NUM> for each engine <NUM>, <NUM> and also includes a power conditioner <NUM> and an energy storage device <NUM>.

The electric generator <NUM> may send electrical power to the power conditioner <NUM>, which may transform, convert, alter, or the like, the electrical energy generated by the generators <NUM> and output a different form, such as by rectifying, inverting, altering voltage or current levels, altering a current frequency, a combination thereof, or the like. The output different form of electrical energy can further be provided or delivered for storage of the energy in the energy storage device <NUM>, or can be provided or delivered to power-consuming systems of the aircraft <NUM>. For the aspect depicted, the electric generators <NUM>, power conditioner <NUM>, and energy storage device <NUM> are all connected by way of an electric communication bus <NUM>, such that the electric generator <NUM> may be electrically connected with the energy storage device <NUM>, a subset thereof, and such that the electric generator <NUM> may provide electrical power to the energy storage device <NUM>. Accordingly, in such an aspect, the system <NUM> may be referred to as a gas-electric propulsion system.

<FIG> depicts a diagram of the system <NUM> in a non-limited example of a dual-spool electrical system <NUM> according to example aspects of the present disclosure. Electrical system <NUM> can include engines <NUM>, <NUM>. Engines <NUM>, <NUM> can be gas turbine engines as described above in <FIG>. Each respective engine <NUM>, <NUM> can include a HP spool <NUM>, <NUM> and a LP spool <NUM>, <NUM>. The HP spool <NUM> of the first engine <NUM> can be configured to drive a first generator <NUM> and provide a first electrical output. The LP spool <NUM> of the first engine <NUM> can be configured to drive a second generator <NUM> and provide a second electrical output. Similarly, the HP spool <NUM> of the second engine <NUM> can be configured to drive a third generator <NUM> and provide a third electrical output. The LP spool <NUM> of the second engine <NUM> can be configured to drive a fourth generator <NUM> and provide a second electrical output.

In an example aspect, the HP spool <NUM> can be used as a starter generator to provide in-air cross-engine electric start. Thus, engines <NUM>, <NUM> can be started from the HP spool <NUM> only. According to another example aspect, the HP <NUM>, <NUM> and LP <NUM>, <NUM> generators on each engine <NUM>, <NUM> can be rated for 250kW and <NUM> kW respectively depending on maximum horsepower extraction capability of the engine. Therefore, each engine <NUM>, <NUM> can provide about 1MW of electric power output through these generators. Each respective electrical output can be connected with a set of converters <NUM>, such as the power conditioner <NUM> of <FIG>. The electrical output from the first generator <NUM> can be connected to a first converter <NUM>, the electrical output from the second generator <NUM> can be connected to a second converter <NUM>. The electrical output from the third generator <NUM> can be connected to a third converter <NUM>. The electrical output from the fourth generator <NUM> can be connected to a fourth converter <NUM>. In one non-limiting example, the set of converters <NUM> can be used to convert AC power generated by the set of generators <NUM>, <NUM>, <NUM>, <NUM> to DC power to be provided to an electrical power distribution bus <NUM>.

A control system <NUM> (e.g., a control system comprising one or more controllers such, as but not limited to, a full authority digital engine controllers (FADEC), a bus power control units (BPCU), a propulsion system controllers (PSC), etc.) can be used to control electrical power flow throughout the aircraft <NUM>. Multiple controllers can be included for redundancy management of power distribution. As shown in <FIG>, the control system <NUM> can include a propulsion system controller (PSC) <NUM> having a controller module <NUM> (shown in FIG. <NUM>) and configured to communicate with the set of converters <NUM>. For example, the propeller 330a, 330b can require 400kW during take-off and 200kW during cruise condition, which can yield substantial engine power available during cruise condition for elective aircraft systems.

According to example aspects of the present disclosure, the PSC <NUM> can also be coupled to the bus power control units (BPCUs) <NUM>, <NUM>. The BPCUs <NUM>, <NUM> can be used to control the distribution of electrical power between multiple distribution buses within the aircraft <NUM>. The BPCUs <NUM>, <NUM> and the PSC <NUM> can command the power resources (e.g., HP and LP generator systems) to provide power to aircraft systems to ensure full flight performance. In other example aspects, the PSC <NUM> or the BPCUs <NUM>, <NUM> can be configured to manage a power split among the first generator <NUM>, the second generator <NUM>, the third generator <NUM>, the fourth generator <NUM>, or an optional battery unit <NUM>. The PSC <NUM> can manage the power split by selectively sending control signals to the set of converters <NUM>. The BPCUs <NUM>, <NUM> can manage the power split by selectively sending control signals to the aircraft systems power distribution bus during specific phases of operation or conditions of operation.

For example, the total engine power extraction is based on the sum of mechanical propulsion power needed by the propellers 330a, 330b and the electrical power needed for the additional aircraft systems. Therefore, if the HP <NUM>, <NUM> and the LP <NUM>, <NUM> are rated for <NUM> kW and <NUM> kW, respectively, then the total electric power output from the engines <NUM> and <NUM> is <NUM> kW through the generators <NUM>, <NUM>, <NUM>, and <NUM>, and propeller requires 200kW power during cruise condition and 400kW during takeoff condition which can be extracted depending on the operational phase of the aircraft <NUM>. Therefore, engine is sized at <NUM> MW to provide maximum electric power extraction of 800kW electrical power from all generators and 200kW propeller power during cruise condition from the engine. During cruise condition, excess power is available for electrical generators and aircraft electrical bus from engine due to reduced propeller power demand. Another non-limiting example can occur if an operational phase of the aircraft <NUM> operates with the first load 382a requiring <NUM> kW, the second load requiring <NUM> kW and the third load requiring <NUM> kW. In response to such load requirements, the PSC <NUM> and the BPCUs <NUM>, <NUM> can manipulate the power distribution to accommodate the operational phase of the aircraft <NUM>.

The PSC <NUM> and the BPCUs <NUM>, <NUM> can also communicate with a vehicle management system (VMS) via the FADEC <NUM>, <NUM> to improve overall flight performance. The FADEC's basic purpose is to provide optimum engine efficiency for the given flight condition. <FIG> shows two BPCUs <NUM>, <NUM> and two FADECs <NUM>, <NUM> for simplification. However, the controllers can be in a variety of different arrangements based on installation, weight impact, etc..

The FADEC <NUM>, <NUM> can receive multiple input variables of the current flight conditions including air density, throttle lever position, engine temperatures, engine pressures, and many other parameters. The inputs are received by an electronic engine controller (EEC) (not shown) located within the FADEC <NUM>, <NUM> and analyzed up to <NUM> times per second. Engine operating parameters such as fuel flow, stator vane position, air bleed valve position, and others can be manipulated in response to the FADEC <NUM>, <NUM> analyzing the multiple input variables. The FADECs <NUM>, <NUM> can be configured to control all aspects of engines <NUM>,<NUM> (e.g., fuel consumption, speed, thrust, etc.). The FADEC <NUM>, <NUM> can also control engine starting and restarting. A benefit of the FADEC <NUM>, <NUM> can include the manufacturer's capability to program engine limitations and receive engine health and maintenance reports. For example, to avoid exceeding a certain engine temperature, the FADEC <NUM>, <NUM> can be programmed to automatically take the necessary measures without requiring pilot intervention.

The electrical system <NUM> can also include the first, second, and third aircraft systems power distribution bus <NUM>. The bus <NUM> can include DC-DC converters <NUM> and DC-AC inverter <NUM>. DC-DC converters <NUM> can be implemented to convert the high voltage (e.g., ±960Vdc) from a hybrid electric bus to the power needs for additional aircraft systems (e.g., 270Vdc). The additional aircraft systems can include a first load 382a, a second load 382b, a third load 382c. While a single respective load 382a, 382b, 382c is shown for brevity, each respective load 382a, 382b, 382c can represent a respective set of electrical loads. The first load 382a, the second load 382b, and the third load 382c can include flight critical systems, non-critical systems, and elective systems.

For example, the flight critical systems can include communication systems, exterior lighting systems, water/waste systems, navigation systems, flight control systems, landing systems, etc. Non-critical systems can include interior lighting systems, refrigeration systems, interior climate control systems, entertainment systems, etc. Elective systems can include environmental systems such as de-icing systems, engine cooling systems, or any other high-power demand system for discretionary operations that require significant electric power coupled to the aircraft <NUM> that can be selectively enabled during particular flight operations, flight parameters, flight demands, or the like. Stated another way, the elective systems of the third load 382c are non-continuously used systems that include a temporary high-power demand, which may otherwise interrupt the ability of the electrical <NUM> to meet the full power demand of the aircraft.

In example aspects, the electrical system <NUM> can include DC-AC inverters <NUM> configured to provide AC power to one or more loads 382a, 382b, 382c on the aircraft, which requires AC power. For instance, DC-AC inverter <NUM> can convert the high voltage DC power from the hybrid electric bus (e.g., ±960Vdc) to a required AC power (e.g., 115Vac, <NUM>-phase, <NUM>) to be delivered to AC loads. There can be multiple DC-DC converters <NUM> and DC-AC <NUM> converters depending on load analysis, redundancy needs and architecture definition. The electrical system <NUM> can also include electrical accumulator unit (EAU) <NUM> to supplement the HP/LP <NUM>/<NUM> generator output power to provide transient performance and an electric start during an engine failure (e.g., act as a micro-grid).

The electrical system <NUM> can also include a battery energy storage system <NUM>. The battery energy storage system <NUM> can be coupled to the electrical power distribution bus <NUM>. The battery energy storage system <NUM> can include a battery energy storage device <NUM> configured to supply a battery power. The power supplied by the battery energy storage system <NUM> can be managed using the PSC <NUM> or the BPCU <NUM>, <NUM> controllers. For example, the PSC <NUM> can be configured to manage the battery power through the power distribution bus <NUM> from the battery energy storage system <NUM> to supplement the electrical power supplied by the first generator <NUM>, the second generator <NUM>, the third generator <NUM>, and the fourth generator <NUM>. In a non-limiting example, the battery <NUM> can be rated for 10kWh. The PSC <NUM> can manage the battery power to use 10kW power for one hour or 100kW (10C) power for six minutes.

<FIG> depicts a flow diagram of an example method <NUM> according to example aspects of the present disclosure. <FIG> depicts steps performed in a particular order for purposes of illustration and discussion. Those of ordinary skill in the art, using the disclosure provided herein, will understand that the method discussed herein can be adapted, rearranged, expanded, omitted, performed simultaneously, or modified in various ways without deviating from the scope of the present disclosure. Method <NUM> can be performed using the control system <NUM>.

At <NUM> the method <NUM> can include splitting total power from the engine <NUM>, <NUM> to a generator <NUM>, <NUM> and a propeller 330a, 330b. The power allocated for the generator <NUM>, <NUM> can be used to manage the first load 382a, the second load 382b, and the third load 382c. The power allocated for the propeller 330a, 330b can be extracted via the low spool <NUM>, <NUM> without compromising the stall margin.

At <NUM> the method <NUM> can include providing a first electrical output from a first generator <NUM> to the electrical power distribution bus <NUM>. The PSC <NUM> can be used or configured to control electrical output from the first generator <NUM>. Converters <NUM>, <NUM>, <NUM>, and <NUM>, or a subset thereof, can be used to convert electrical output from the first generator <NUM> into power suitable for the electrical power distribution bus <NUM> (e.g., convert AC power from the first generator <NUM> to DC power).

At <NUM> the method <NUM> the PSC <NUM> can be used to control electrical power from the second generator <NUM> to the electrical power distribution bus <NUM>.

At <NUM> the method <NUM> the PSC <NUM> can be used to couple a first aircraft systems bus, a second aircraft systems bus, and a third aircraft systems bus <NUM> to the power distribution bus <NUM>. In example aspects, BPCUs <NUM>, <NUM> can be used to control power to the secondary aircraft systems bus <NUM>. The secondary aircraft systems bus <NUM> can include DC-DC converters <NUM> and DC-AC converters <NUM>. Secondary aircraft systems can include flight controls, galley loads, and electrical actuators. Secondary aircraft systems can also include flight critical and non-critical loads <NUM>. As shown in <FIG>, BPCUs <NUM>, <NUM> can be used to manage the power allocation to secondary aircraft systems.

At <NUM> the method <NUM> can include managing, by a control system <NUM>, a power allocation among the first generator <NUM>, the second generator <NUM>, the propulsion system <NUM>, the first aircraft systems bus, the second aircraft systems bus, and the third aircraft systems bus <NUM>. According to example aspects, the electrical system <NUM> can leverage the available multi-spool power extraction to improve energy sources for the secondary aircraft systems. The power allocation can be achieved by intelligent controllers such as PSC <NUM> and BPCUs <NUM>, <NUM>. The PSC <NUM> and BPCUs <NUM>, <NUM> can interface with FADECs <NUM>, <NUM>. These controllers can work together to command the power sources (e.g., HP and LP generator systems) to provide the required power for propulsion and secondary aircraft systems. For example, PSC <NUM> can be used to allocate power among the first generator <NUM>, the second generator <NUM>, the third generator <NUM>, the fourth generator <NUM>. In other example aspects, the PSC <NUM> and BPCUs <NUM>, <NUM> can be used to allocate power among additional aircraft systems through the FADECs <NUM>, <NUM>. For instance, the power allocation can be managed depending on phase of the flight envelope (e.g. taxi, take off, climb, cruise, descent, land, etc.). In other example aspects, the power allocation can be determined depending on aircraft load demand and depending on emergency conditions (e.g., single engine failure). This method of power allocation can be changed at any time during the flight phase based on a number of conditions. This method of power allocation provides an efficient usage of available power instead of dedicated separate power sources for secondary systems. Management of electrical power from HP <NUM> and LP <NUM> spool of each engine and additional systems <NUM> can help to improve full flight performance while achieving greater fuel efficiency. For example, if <NUM> MW total power is available from both engines, then the power can be spilt <NUM>% for cruise propulsion operation, <NUM>% for flight critical loads, and <NUM>% for non-critical loads 382b, and <NUM>% for elective loads 382c. The total power split can also change at any time during the current flight phase or future flight phases. A benefit of splitting the power can be to provide optimum usage of available power instead of dedicated separate power sources.

<FIG> depicts an example controller module <NUM> of the control system <NUM> according to aspects of the present disclosure. The controller module <NUM> can include one or more processors(s) <NUM> and one or more memory devices(s) <NUM>. The one or more processors(s) <NUM> can include any suitable processing device, such as a microprocessor, micro-control device, integrated circuit, logic device, or the like. The one or more memory devices(s) <NUM> can include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, RAM, ROM, hard drives, flash drives, or the like.

The one or more memory device(s) <NUM> can store information accessible by the one or more processor(s) <NUM>, including computer-readable instructions that can be executed by the one or more processors(s) <NUM>. The instructions can be any set of instructions that when executed by the one or more processors(s) <NUM>, cause the one or more processors(s) <NUM> to perform operations. In some aspects, the instructions can be executed by the one or more processor(s) <NUM>, to cause the one or more processors(s) <NUM> to perform operations. In some aspects, the instructions can be executed by the one or more processor(s) <NUM> to cause the one or more processor(s) to perform operations, such as any of the operations and functions for which the control module <NUM> is configured. For instance, the operations can be used for performing method <NUM>, as described herein, or any other operations or functions of the one or more control system. The instructions can be software written in any suitable programming language or can be implemented in hardware. Additionally, or alternatively, the instructions can be executed in logically or virtually separate threads on processor(s) <NUM>. The memory device(s) <NUM> can further store data that can be accessed by the processor(s) <NUM>. For example, the data can include data indicative of power flows, current flows, temperatures, actual voltages, nominal voltages, gating commands, switching patterns, or any other data/or information described herein.

The controller module <NUM> can also include a communication interface <NUM>. The communication interface <NUM> can include suitable components for interfacing with one or more network(s), devices, or the like, including for example, transmitters, receivers, ports, control devices, antennas, or other suitable components. For example, the communication interface <NUM> can be configured to communicate with a control system, such as the control module <NUM>.

In this way, example aspects of the present disclosure can provide a number of technical effects and benefits. For instance, the technical effect of above described aspects enable managing a power split between the multiple high power demand aircraft systems and can help improve power utilization during full flight performance. Improved power utilization can result in, for example, achieving better fuel efficiency. Managing the power split can also eliminate the need for larger engines with dedicated power sources for high demand power-consuming aircraft systems, such as, but not limited to, propulsion systems, non-propulsions systems, environmental control systems, additional customize high power demand systems. Another benefit of the present disclosure can include eliminating the need for separate power generators coupled to an engine for powering flight critical systems, non-critical systems, and elective systems. The system also eliminates the need to have separate gearboxes or gearbox pads to accommodate additional generators or power sources dedicated to specific aircraft systems.

Another benefit of the present disclosure can include the multi-spool power extraction from the engine, which can enable full-flight envelope performance by providing sufficient engine stall margin during corner point operating conditions of the flight envelope. Also, by utilizing higher voltage power (e.g. power greater than <NUM> VDC or <NUM> VAC, the power distribution feeder cable weight can be reduced for the power bus for additional systems.

Another benefit of the present disclosure can include optimized aircraft flight performance by balancing fuel efficiency and full flight operability due to electrical communicate between the FADEC <NUM>, <NUM>, the BPCU <NUM>, <NUM>, and the PSC controller <NUM>. Also, the optional battery <NUM> can provide hybrid electric propulsion capability then it can be used for emergency conditions (e.g. engine failure/s) to provide power for flight controls and essential secondary loads to land safely. This will eliminate need for separate emergency power source e.g. RAT or APU.

Many other possible configurations in addition to those shown in the above figures are contemplated by the present disclosure. To the extent not already described, the different features and structures of the various aspects can be used in combination with others as desired. That one feature cannot be illustrated in all of the aspects is not meant to be construed that it cannot be, but is done for brevity of description. Thus, the various features of the different aspects can be mixed and matched as desired to form new aspects, whether or not the new aspects are expressly described. Combinations or permutations of features described herein are covered by this disclosure.

Claim 1:
An electrical system for an aircraft (<NUM>) having a first turbine engine (<NUM>), the first turbine engine having a first high pressure (HP) spool (<NUM>) and a first low pressure (LP) spool (<NUM>), the electrical system comprising:
a first generator (<NUM>) coupled to the first HP spool (<NUM>) driven by the first turbine engine (<NUM>) and providing a first electrical output;
a second generator (<NUM>) coupled to the first LP spool (<NUM>) driven by the first turbine engine (<NUM>) and providing a second electrical output;
the first generator (<NUM>) and the second generator (<NUM>) coupled to an electrical power distribution bus (<NUM>) that provides electrical power to a first load (382a), a second load (382b), and a third load (382c);
a propulsion system controller (<NUM>) coupled to the electrical power distribution bus (<NUM>);
a first aircraft systems bus (<NUM>), a second aircraft systems bus (<NUM>), and a third aircraft systems bus (<NUM>) coupled to the electrical power distribution bus (<NUM>);
a first load (382a), a second load (382b), and a third load (382c) coupled to the first aircraft systems bus (<NUM>), the second aircraft systems bus (<NUM>), and the third aircraft systems bus (<NUM>) respectively; and
a control system (<NUM>) configured to allocate power from the first generator (<NUM>) and the second generator (<NUM>) via the propulsion system controller (<NUM>) to at least one of the first load (382a) via the first aircraft systems bus (<NUM>), the second load (382b) via the second aircraft systems bus (<NUM>), and the third load (382c) via the third aircraft systems bus (<NUM>).