Aircraft Range Extension Energy Pod (REEP)

Abstract

The energy source for the aircraft includes an enclosure, an engine, a generator, at least one fuel tank configured to supply fuel to the engine, and an electrical connector for outputting electrical power generated by the generator to at least one electrical component or electrical bus of the aircraft, wherein the engine, the generator, and the at least one fuel tank are each contained within the enclosure.

Description

[Background technology]

[0001] CROSS-REFERENCE TO RELATED PATENT APPLICATIONS This application claims the benefit of U.S. Provisional Patent Application No. 63 / 280,615, filed November 17, 2021, U.S. Provisional Patent Application No. 63 / 163,165, filed March 19, 2021, and U.S. Provisional Patent Application No. 63 / 151,760, filed February 21, 2021, the entire contents of each of which are incorporated herein by reference in their entirety.

[0002] This application relates to an aircraft Range Extending Energy Pod (REEP).

[0003] Various types of aircraft may be used to transport goods or people, or for recreational use. Various types of aircraft may be specified with a range based on fuel capacity, type of aircraft, type of engine or other propulsion system, weather conditions, etc. The range may indicate how far the aircraft can be safely flown under given conditions. Attempting to fly an aircraft beyond a particular range may result in the aircraft not operating safely. Summary of the Invention

[0004] In one embodiment, an energy source for an aircraft includes an enclosure, an engine, a generator, at least one fuel tank configured to supply fuel to the engine, and an electrical connector for outputting electrical power generated by the generator to at least one electrical component or electrical bus of the aircraft, wherein the engine, the generator, and the at least one fuel tank are each contained within the enclosure.

[0005] In one embodiment, a method for using a removable energy source on an aircraft includes installing the removable energy source on the aircraft. The removable energy source includes an engine and a generator, the engine and the generator each housed in an enclosure. The method further includes connecting a first electrical connector of the removable energy source to a second electrical connector of the aircraft. The method further includes outputting electrical power from the generator of the removable energy source to at least one electrical component or electrical bus of the aircraft.

[0006] In one embodiment, the energy source for the aircraft includes an enclosure, an engine, a generator, and mounting hardware for mounting the energy source to the aircraft. The energy source for the aircraft further includes an electrical connector for outputting electrical power generated by the generator to at least one electrical component or electrical bus of the aircraft. The engine and the generator are housed within the enclosure. [Brief description of the drawings]

[0007] [Figure 1] FIG. 1 is a perspective view of an aircraft having a range extension energy pod (REEP) in accordance with an illustrative embodiment. [Diagram 2] FIG. 1 is a perspective view of an exemplary range-extended energy pod (REEP) in accordance with an exemplary embodiment. [Diagram 3] FIG. 6 is a side view of the REEP of FIG. 5 in accordance with an exemplary embodiment. [Figure 4] FIG. 6 is a front view of the REEP of FIG. 5 in accordance with an exemplary embodiment. [Diagram 5] 6 is a perspective view of the REEP of FIG. 5 according to an exemplary embodiment, showing that the enclosure of the REEP is partially transparent. [Figure 6] FIG. 2 is a perspective view of another range-extended energy pod (REEP) in accordance with an illustrative embodiment, showing that the REEP enclosure is partially transparent. [Figure 7] FIG. 7 is a front view of the REEP of FIG. 6 in accordance with an exemplary embodiment. [Figure 8A] FIG. 7 is a side view of the REEP of FIG. 6 in accordance with an exemplary embodiment. [Figure 8B] 4 is a flowchart illustrating an example method of using REEP, according to an example embodiment. [Figure 9A] FIG. 1 illustrates an example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 9B] FIG. 1 illustrates an additional example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 10A] FIG. 1 is a block diagram illustrating a first aircraft control system for use with the flexible architecture of an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 10B] FIG. 1 is a block diagram illustrating a second aircraft control system for use with the flexible architecture of the aerospace hybrid system in accordance with an illustrative embodiment. [Figure 11] FIG. 1 illustrates a first example of an aircraft that can use the flexible architecture of an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 12] FIG. 1 illustrates a second example of an aircraft that can use the flexible architecture of an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 13] FIG. 10 illustrates a third example of an aircraft that may use the flexible architecture of an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 14] 1 is a flowchart illustrating a first exemplary method for using a flexible architecture of an aerospace hybrid system during different flight phases of an aircraft equipped with a main propulsion propeller, in accordance with an exemplary embodiment. [Figure 15] 1 is a flowchart illustrating a second exemplary method for using a flexible architecture of an aerospace hybrid system during different flight phases of an aircraft equipped with a main propulsion propeller, in accordance with an exemplary embodiment. [Figure 16A]FIG. 1 illustrates an example of a flexible architecture for an aerospace hybrid system having a flywheel in accordance with an illustrative embodiment. [Figure 16B] FIG. 1 illustrates an example of a flexible architecture for an aerospace hybrid system having a flywheel and a spring coupler in accordance with an illustrative embodiment. [Figure 17] FIG. 1 illustrates a perspective view of an example flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 18] FIG. 18 is a top view of the example flexible architecture of FIG. 17 in accordance with an exemplary embodiment. [Figure 19] FIG. 18 is a side view of the example flexible architecture of FIG. 17 in accordance with an exemplary embodiment. [Figure 20] FIG. 1 is a perspective view of another example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. [Figure 21] FIG. 1 illustrates an example of downstream and upstream components for propelling an aircraft in accordance with an illustrative embodiment. [Figure 22] 1 is a schematic diagram of an example cooling system for a hybrid power plant according to an exemplary embodiment. [Figure 23] FIG. 1 illustrates an example of a hybrid power plant with a cooling system according to an exemplary embodiment. [Figure 24] 24 is a cross-sectional view of an example hybrid power plant with the cooling system of FIG. 23 according to an exemplary embodiment. [Diagram 25] 24 is a partial cross-sectional perspective view of an example hybrid power plant with the cooling system of FIG. 23 according to an exemplary embodiment. [Figure 26] 2 is a schematic diagram of a second example cooling system for a hybrid power plant in accordance with an illustrative embodiment. [Figure 27] 1 is a schematic diagram of a third example cooling system for a hybrid power plant in accordance with an illustrative embodiment. [Figure 28] 11 is a schematic diagram of a fourth example cooling system for a hybrid power plant in accordance with an illustrative embodiment. [Figure 29] 13 is a schematic diagram of a fifth example cooling system for a hybrid power plant in accordance with an illustrative embodiment. [Diagram 30] 1 is a top view of an example hybrid power plant with a cooling system according to an exemplary embodiment; [Diagram 31] 30A is a cross-sectional view taken along line AA of FIG. 30, illustrating an example of the hybrid power plant of FIG. 30, according to an exemplary embodiment. [Diagram 32] 31 along line BB of FIG. 31, illustrating an example of the hybrid power plant of FIG. 30, according to an exemplary embodiment. [Diagram 33] 31 is an alternative view of the example hybrid power plant of FIG. 30 showing engine cooling fins details in accordance with an exemplary embodiment. [Diagram 34] 31 is a side view of the example hybrid power plant of FIG. 30 with a cooling system according to an exemplary embodiment. [Diagram 35] 1 is a diagram of an example system for providing a stable voltage to a direct current (DC) bus in accordance with an illustrative embodiment. [Diagram 36] 1 is a flowchart illustrating an example method for maintaining a stable DC bus voltage based on communications from an aircraft-level controller in accordance with an example embodiment. [Figure 37] 4 is a flowchart illustrating an example method for maintaining a stable DC bus voltage based on measurements by a hybrid generator-level controller in accordance with an example embodiment. [Figure 38] 1 is a diagram of an example computing environment in accordance with an illustrative embodiment. [Figure 39] FIG. 2 is a cross-sectional side view of an enclosure having noise reduction components according to an exemplary embodiment. [Figure 40A] FIG. 2 is a front perspective view illustrating an air inlet of an exemplary enclosure having noise reducing components therein, according to an exemplary embodiment. [Figure 40B] FIG. 40B is a side view of the exemplary enclosure of FIG. 40A according to an exemplary embodiment. [Figure 40C]FIG. 40B is a rear view of the exemplary enclosure of FIG. 40A according to an exemplary embodiment. [Figure 40D] FIG. 40B is a top view of the exemplary enclosure of FIG. 40A according to an exemplary embodiment. [Diagram 41] FIG. 40B is a rear perspective view showing an exhaust vent of the enclosure of FIG. 40A according to an exemplary embodiment. [Diagram 42] FIG. 40B is a top perspective view of a noise-reducing channel within the noise-reducing chamber of the enclosure of FIG. 40A according to an exemplary embodiment. [Diagram 43] FIG. 2 is a top perspective view of another exemplary enclosure having noise reducing components therein, in accordance with an exemplary embodiment. [Diagram 44] FIG. 44 is a side view of the enclosure of FIG. 43 in accordance with an exemplary embodiment. [Diagram 45] FIG. 44 is a front view of the enclosure of FIG. 43 according to an exemplary embodiment. [Figure 46] FIG. 44 is a perspective view of the enclosure of FIG. 43 according to an exemplary embodiment, showing that the enclosure is partially transparent. [Figure 47] 1 is a perspective view of another example of an enclosure according to an exemplary embodiment, showing that the enclosure is partially transparent and has noise reducing components therein. [Figure 48] FIG. 48 is a top view of the enclosure of FIG. 47 in accordance with an exemplary embodiment. [Figure 49] FIG. 48 is a side view of the enclosure of FIG. 47 according to an exemplary embodiment. [Figure 50] FIG. 48 is a rear view of the enclosure of FIG. 47 according to an exemplary embodiment. [Figure 51] FIG. 48 is a rear view of the enclosure similar to FIG. 47, except that the enclosure is shown as opaque, according to an exemplary embodiment. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0008] Today's aviation industry is undergoing a revolution with the widespread adoption of electric propulsion. Many vehicles around the world are under development in which power to the fans / propellers / rotors used for propulsion, lift and control is provided via electrical wires rather than a mechanical shaft. The use of electricity to transmit power remotely from its generation and storage devices is a design factor for many new designs.

[0009] Electric propulsion depends on three key factors: system voltage, energy, and power. Energy is the total storage capacity (measured in kilowatt hours or kWh) and power is a measure of the flow of energy (measured in kW). Additionally, many new generation aircraft using distributed electric propulsion may rely on direct current (DC) for power and energy storage and transmission.

[0010] A common device for storing electrical energy and providing power is the battery pack. In the case of batteries, the conversion of energy involves chemistry. Energy or power is added to the battery pack for storage (resulting in a chemical change) and then extracted from the battery pack in a reverse reaction for use when needed. Another device used for energy or power storage is the supercapacitor or ultracapacitor. Although these devices can deliver very high power levels, they are rarely chosen as the primary energy storage device for aircraft because the total energy storage is very low for a given product weight.

[0011] Battery packs may consist of individual cells arranged in modules, and further arranged in battery packs consisting of multiple battery modules or packs. Safety regulations and performance requirements may require the use of battery management systems (to maintain balanced cell voltages through multiple charge / discharge cycles), cooling systems, ventilation in case of fire or unwanted release of other chemicals / gases, and / or safety circuits. All of these support systems may contribute to the total mass of the battery pack. In many cases, the specific energy (measured in watt-hours per kilogram (Wh / kg)) of a complete battery pack suitable for safe operation on a manned aircraft will be much lower than the energy that may be required for the aircraft design and a given mission of the aircraft. That is, current batteries may be too heavy for many of today's aircraft designs and / or many missions / flight plans desired to be performed on today's aircraft.

[0012] Described herein are hybrid-electric gensets that help address the problem of low specific energy of batteries. In particular, the high energy density of liquid fuel can be used by an engine to convert shaft power, which can then be converted to electricity using a generator. In this way, depending on the amount of fuel carried on board the aircraft, the specific energy of the entire system can be much greater than that of a battery. As described herein, the hybrid-electric gensets of various embodiments of the present application can provide more than six times (6X) the specific energy (energy per unit mass) of a battery at a desired power level.

[0013] These hybrid generators may include components for physical mounting, wiring, liquid fuel storage, and airflow interface and interaction between the aircraft and the hybrid generator. In some cases, an aircraft may be completed with no design provisions for the addition of an internal generator, and only plans for battery storage of energy. Therefore, in certain situations, it may be desirable to have supplemental energy or power without redesigning major portions of the aircraft. Therefore, the availability of a Range Extension Energy Pod (REEP) would be very helpful. The REEP described herein can perform the exact same functions as an external battery: mounting physical items and electrical connections.

[0014] Thus, various embodiments of an aircraft range extension energy pod (REEP) are described herein. The REEP can include a hybrid power plant and / or various other desired elements in a single compact package or enclosure. The enclosure can include an engine, a generator, one or more fuel tanks, etc., and anything else for powering an electric aircraft. The REEP thus has the advantage that it can be removably connected to an aircraft whenever additional power is needed. For example, some aircraft may have a certain limited range. However, the REEP described herein can be attached to such an aircraft and provide additional power, thereby extending the range of the aircraft. In this manner, the REEP can function similarly to a battery from the aircraft's perspective in that it can simply be plugged into the aircraft's existing electrical system via a high voltage bus or the like and function as a battery to power the aircraft.

[0015] The REEP may also include mounting hardware so that the REEP can be easily mechanically secured to the aircraft and removed from the aircraft as desired. For example, the REEP may be bolted to the aircraft and an electrical connector, such as two high voltage wires, may be plugged in. The aircraft may then be provided with a large amount of power or energy beyond what a permanently attached on-board system could provide. The power or energy from the REEP may be used to drive a propulsion system (e.g., an electric motor configured to rotate a rotor, propeller, etc.) and / or may be used for other purposes by the aircraft, such as powering other electronics (e.g., accessories) or charging the aircraft's batteries. Thus, the embodiments described herein advantageously provide additional power and energy for use by the aircraft.

[0016] Battery-powered electric aircraft may be able to perform missions or flights that conventional aircraft using conventional power plant solutions cannot or would not be permitted to perform. For example, electric aircraft may be able to take off or land in tight spaces where conventionally powered aircraft cannot take off or land. Electric aircraft may also be permitted to operate in areas where conventional aircraft are not permitted to operate due, for example, to the large amount of noise generated by some conventional aircraft.

[0017] However, batteries can be very heavy and often provide insufficient energy or power to enable a particular mission or flight path (e.g., they may not provide enough energy or power for certain power-intensive tasks such as takeoff or landing, or may not provide enough energy for a flight route of sufficient length as needed). As such, various embodiments of hybrid electric generators are described herein that convert liquid fuel into electrical energy and power that can be utilized by an electric aircraft (or any aircraft with electrical components). The range extension energy pods (REEPs) described herein may advantageously include any fuel, power conversion, heat treatment, and wiring to allow the REEP (which may function as a hybrid battery) to be completely independent of the aircraft body, simplifying installation. By way of example only, the REEP may be attached to the aircraft using four bolts for support and two wires for transmitting electricity. As such, the REEPs described herein may be attached to aircraft that have their own propulsion and power generation equipment, and those aircraft may fly with or without the REEP attached. Alternatively, the REEP can provide additional power beyond the power the aircraft would have alone without the REEP, for example to extend the range of the aircraft, the speed at which the aircraft can fly, etc.

[0018] By way of example only, an embodiment may include storage of a quantity of fuel suitable to provide 185 kilowatts (kW) of power at 800 volts direct current (VDC) nonstop for three hours. This corresponds to approximately 555 kilowatt-hours (kWh) of energy. Such an embodiment may have a weight of approximately 450 kg and provide an energy density of over 1200 watt-hours per kilogram (Wh / kg). Current battery packs currently available for the aviation industry can only provide a maximum energy density of approximately 200 Wh / kg when considered at the pack level, including cooling components and necessary battery management hardware to maintain safe operation. Thus, the exemplary embodiments described herein may provide at least a six-fold benefit based on energy per unit weight compared to battery systems. The embodiments herein also provide a significant simplification of use for a significant increase in energy density, as the REEP described herein may be removably attached to an aircraft using, for example, only four bolts and simple electrical connections.

[0019] An exemplary embodiment of the REEP may include an integrated hybrid generator (e.g., any of the flexible architectures described herein, such as under the Flexible Architecture Elements heading below), which may include provision for power conversion from liquid fuel to direct current (DC) power (e.g., including any of the components described herein in connection with power components, such as under the Direct Current (DC) Bus Elements heading below). The hybrid generator may further include one or more integrated cooling systems (e.g., any of the cooling systems or elements described herein, such as under the Air Cooling Elements heading below). An exemplary embodiment of the REEP may further include one or more storage tanks for liquid fuel and components for safely connecting the stored fuel to the hybrid generator. An exemplary embodiment of the REEP may further include noise reduction elements (e.g., any of the noise reduction components described herein, such as under the Noise Reduction Elements heading below).

[0020] Exemplary embodiments of the REEP may further include various physical structures such as enclosures (e.g., cowlings) for the elements of the REEP, mounting hardware such as physical frames for the elements of the REEP, holes for bolts used to mount the REEP, etc. For example, the structural frame of the REEP may be attached to the aircraft with as few as four mounting points. In other embodiments, a different number of mounting points may be used. The mounting points may be located at or extend below the bottom of the REEP (e.g., when mounting above an existing aircraft surface), located on or extend above the system (e.g., when mounting below an existing aircraft surface), or configured in any other manner necessary to provide mounting points for mounting the REEP to the aircraft. Mounting hardware may also include any aspect of the REEP designed to facilitate connection of the REEP to the aircraft. For example, any type of mechanical structure configured to mount to the aircraft may be part of the mounting hardware. For example, a cowling, housing, or enclosure of the REEP may be mounting hardware if the cowling, housing, or enclosure is designed to be flush with and welded or otherwise secured to a surface or portion of the aircraft. If the frame to which the REEP components within a cowling, housing, or enclosure are attached is also used to securely attach the REEP to the aircraft (e.g., by bolting a portion of the REEP frame to a portion of the aircraft), then the frame or structural components of the REEP may also be part of the mounting hardware.

[0021] The REEP enclosure may be or include an aerodynamic firewall package to control cooling flow, limit transmission of system noise, reduce aerodynamic drag, and / or provide a clean integrated package. In other words, the enclosure may house the elements of the REEP making them more aerodynamic, more visually appealing, safer, and less noisy. The entire REEP including the enclosure may be mounted to the aircraft using, for example, four bolts. The mounting points may be spaced apart with sufficient lateral and vertical extent to provide adequate support for the hybrid generator in terms of strength and rigidity. In one embodiment, the total weight of the exemplary REEP is approximately 1000 pounds (lbs), so in such an embodiment, the mounting hardware may include 4 x AN-4 bolts (1 / 4-28).

[0022] Electrical connectors can be used to transfer high voltage current (energy and power). Depending on the voltage, this can be a single pair of wires (positive and negative) or multiple pairs of wires. For example, one embodiment of the REEP can deliver up to 185 kW of power at 800VDC. Thus, such an embodiment can have a DC current of 230 Amps (A) and can use a single pair of wires using a 3 / 0 or 4 / 0 wire size. Various embodiments may use other sizes of wires and numbers of wires.

[0023] Various embodiments may include electrical or electronic communication between the REEP and the aircraft's control systems, allowing the aircraft systems to affect and / or control activation or deactivation of the REEP and / or the flow of energy and power from the REEP. This same communication interface may provide system health and stability information of the REEP for use by the aircraft and pilot. As such, additional wiring may be used to connect the REEP to the aircraft for communication and control purposes.

[0024] Various embodiments also provide REEPs with good aerodynamic performance. Because the REEPs are configured to be attached to an aircraft, such as to the exterior surface of the aircraft, the REEPs advantageously have an aerodynamic profile that does not adversely affect the flight of the aircraft. In other embodiments, the REEPs may fit into an aircraft battery enclosure within the aircraft fuselage or other aerodynamically designed portion of the aircraft, such that the enclosure does not have to be aerodynamically designed. Examples of aerodynamic shapes include enclosures with rounded exterior shells to provide low drag when connected to the aircraft. Such shapes (e.g., as shown in Figures 6-8) can provide surfaces that are separate from adjacent aircraft surfaces, so that the REEPs are designed to be aerodynamic in their own right. Other embodiments (e.g., as shown in Figures 1-5) may have enclosures with exterior shells designed to blend in and coordinate with adjacent aircraft surfaces. To facilitate such blending and aerodynamic tuning, the REEP is designed with a primary exterior shell that comprises up to 80% of the shell, and the remaining up to 20% of the shell is designed for a specific aircraft application to provide aerodynamic margins specific to each aircraft. In various embodiments, other ratios other than 80 / 20 may be used for the set shell / enclosure portion and the customizable shell / enclosure portion based on the target aircraft. Thus, the REEP can be optimized to be aerodynamic on a variety of aircraft.

[0025] In various embodiments, the REEPs described herein may be used to retrofit aircraft that may not have been designed to have the interior space or other configuration / space to hold a hybrid power plant therein. Thus, an aircraft that does not have a hybrid power plant may have one or more hybrid power plants added in the form of a REEP described herein. In other words, the REEPs described herein may be removable (e.g., can be attached and detached from the aircraft for a particular mission / application) or may be used to retrofit an aircraft (e.g., even an aircraft designed without provisions for hybridization may be converted or modified to include a hybrid power plant if a more permanent use of the REEP is desired for the aircraft).

[0026] FIG. 1 illustrates a perspective view of an aircraft 10100 having a range extension energy pod (REEP) 10115, according to an exemplary embodiment. The REEP 10115 is mounted centrally between the wing 10110 and the fuselage 10105 of the aircraft 10100. It may be mounted in such a manner so that the REEP 10115 does not upset the balance of the aircraft 10100. In various embodiments, more than one REEP may be used. In such embodiments, the multiple REEPs may each be positioned along the central axis of the aircraft, or may be positioned equidistant from the central axis such that the balance of the aircraft is maintained even when multiple REEPs are used. In embodiments where the REEPs are positioned on either side of the central axis of the aircraft, it may be desirable to have an even number of REEPs to balance the aircraft. In various embodiments, the REEPs may be mounted, for example, on the underside or upper surface of the wing away from the fuselage. As such, the REEPs may be connected to the aircraft in a manner similar to drop tanks that may be attached to the aircraft at specific hard points on one or more wings (e.g., to be suspended from the wing). In various embodiments, the REEPs may be connected to other portions of the aircraft as needed. As described herein, the REEP 10115 may be removable from the aircraft 10100 so that it is used only when needed (e.g., for longer flights). The REEP may connect to the aircraft 10100 via mechanical fasteners, such as bolts, or may electrically connect to the aircraft 10100 via a wired connector that may be removably attached to a connector on the aircraft 10100. Although not shown in FIG. 1, the aircraft 10100 may have a propulsion system that is powered by electricity (e.g., DC power from the REEP 10115) such that the REEP at least partially powers the aircraft 10100.

[0027] In various embodiments, the REEP may be connected to the aircraft in a more permanent or non-removable manner. For example, the housing of the REEP may be connected to a portion of the aircraft using a more permanent fastening method, such as welding, riveting, or bonding the housing to a wing, fuselage, or other component of the aircraft. In this manner, the REEP is more permanently attached to the aircraft, for example, with a more permanent attachment than bolts. Thus, the described REEP may be a self-contained engine, generator, fuel tank, firewall, and noise abatement device, all contained within a housing or enclosure that may be removable from the aircraft by disconnecting mechanical connectors, or may be more permanently attached to the aircraft. In either case, the enclosure / housing may wall off the internal components at a separate location outside the fuselage, thereby advantageously isolating these components from the fuselage and providing risk management benefits (e.g., less risk of fire at or near the fuselage). In various embodiments, the REEP accomplishes these and other objectives by having various components within a housing or enclosure, with the only components that may pass through the housing or enclosure being wiring for power output and / or control signals. The enclosure or housing may be an aerodynamic enclosure as described and illustrated herein, such that the REEP does not create significant drag on the exterior of the fuselage and may be removable as described herein, or at least may be placed on the aircraft in a more permanent state without redesigning an existing aircraft to accommodate the REEP. For example, bolts, nuts, and other hardware that may be used to connect the REEP to the aircraft may be considered removable because they may be non-destructively loosened (e.g., may be used to attach the REEP to the aircraft multiple times). In various embodiments, when components of the REEP are welded, riveted, bonded, etc. to the aircraft, those fastening mechanisms may only be destructively removed, and may be used when a more permanent connection of the REEP to the aircraft is desired.In various embodiments, a non-destructively removable mechanism may be used alone, a destructively removable mechanism may be used alone, or both destructively and non-destructively removable mechanisms may be used to fasten or otherwise secure the REEP to the aircraft. Additionally, the enclosure of the REEP described herein also functions as a mechanism to transmit and reduce noise output by the REEP (e.g., including the noise reduction components described herein). As such, the REEP advantageously provides a package of components to removably or permanently supply power to an aircraft without the need to redesign the aircraft or add components such as noise reduction components beyond the installation of the REEP itself.

[0028] 2 is a perspective view of an example of a range extension energy pod (REEP) 10200 in accordance with an exemplary embodiment. The REEP 10200 includes an enclosure 10202 and an air intake 10204. The air intake 10204 may be used to intake air for engines, cooling, etc. As described herein, the REEP 10200 may be aerodynamically designed to be mounted on an aircraft, such as the aircraft 10100 of FIG.

[0029] FIG. 3 is a side view of the REEP 10200 of FIG. 2 according to an exemplary embodiment. FIG. 4 is a front view of the REEP 10200 of FIG. 2 according to an exemplary embodiment. The REEP 10200 further includes mounting hardware 10206 and wiring 10208 so that the REEP 10200 can be mounted to an aircraft, such as the aircraft 10100 described herein. For example, the mounting hardware 10206 can be connected to a structural frame that supports the components of the REEP 10200, to which the components of the REEP 10200 can be mounted. The mounting hardware 10206 can further accommodate fasteners, such as bolts, for mounting the REEP 10200 to the aircraft. The wiring 10208 can output power to the aircraft or other power consuming or distributing devices. The wiring 10208 may output AC power generated by an electric machine (e.g., a generator) or may output DC power (e.g., after conversion from AC power to DC power by an inverter of the REEP 10200).

[0030] FIG. 5 is a perspective view of the REEP 10200 of FIG. 2 according to an exemplary embodiment, showing that the enclosure 10202 of the REEP is partially transparent. The REEP 10200 shown in FIG. 5 further includes a hybrid generator as described herein (e.g., flexible architecture elements, air cooling elements, direct current (DC) bus elements, and / or noise reduction elements paragraphs below). As such, the enclosure 10202 may include, for example, an engine, a generator, a cooling system, noise reduction elements, power electronics for providing DC power and energy, etc. Additionally, a fuel tank 10504 may be included to store fuel consumed by the engine of the hybrid generator. Although only a single fuel tank 10504 is shown, another fuel tank may be located on the opposite side of the enclosure 10202 to balance the weight of the REEP 10200. Fuel from the fuel tank 10504 may be transferred to the engine of the REEP 10200 via tubes 10506, for example. Also, although not shown in FIG. 5, one or more of the fuel tanks (e.g., fuel tank 10504) may have openings or ports that may be in or through the enclosure 10202 to facilitate filling the fuel tank. Such openings may also be fitted with caps or other similar covers to prevent fuel from leaking from the fuel tank while the fuel tank is not filled. In various embodiments where multiple fuel tanks are present, balance tubes may be provided that extend between the multiple fuel tanks to balance fuel among the multiple fuel tanks and facilitate filling multiple fuel tanks at one time.

[0031] FIG. 6 is a perspective view of another range-extended energy pod (REEP) 10600, according to an exemplary embodiment, showing that the REEP's enclosure 10602 is partially transparent. FIG. 7 is a front view of the REEP 10600 of FIG. 6, according to an exemplary embodiment. FIG. 8A is a side view of the REEP 10600 of FIG. 6, according to an exemplary embodiment. The enclosure 10602 may further house a hybrid generator 10608 and a fuel tank 10610, similar to FIG. 5 described above. The REEP 10600 may also include a structural frame 10620 to which the components of the REEP 10600 may be permanently attached and may provide for removable attachment to an aircraft (e.g., the structural frame 10620 may also function as mounting hardware).

[0032] Air can enter the cooling system 10618 at an inlet 10624, which can be formed in an interior wall 10626 that separates the main compartment of the enclosure 10602 from the noise reduction chamber 10604 of the enclosure 10602. Similarly, another noise reduction chamber 10606 can be attached to the rear of the enclosure 10602 and can further have a wall that separates the main compartment of the enclosure 10602 from the noise reduction chamber 10604. The noise reduction chamber 10604 and / or 10606 can have noise reduction elements therein (e.g., channels formed by walls of a noise attenuating material as described herein). In this manner, the REEP 10600 can be designed to minimize the noise output by the REEP 10600 during use.

[0033] The REEP 10600 further includes wiring 10612 configured to removably electrically connect the REEP 10600 to an aircraft, as described herein. The power and energy provided to the aircraft may be DC power and energy, such as provided by a battery pack. In this manner, an aircraft designed to operate on battery power can be powered by the REEP 10600 without modifying the manner in which the battery-powered components of the aircraft operate.

[0034] The REEP 10600 further includes an air inlet filter 10616 that introduces filtered air to the engine 10608 for combustion and power generation.

[0035] 8A further shows that the noise reduction chamber 10606 of the enclosure 10602 can have an exhaust vent 10602. In this way, the air used by the cooling system 10618 can have a way to exit, but the air passes through the noise reduction chamber to reduce the noise emitted into the atmosphere.

[0036] FIG. 8B is a flow chart illustrating an example method 10900 of using a REEP. In operation 10902, a removable energy source or REEP is attached to an aircraft. As described herein, the REEP can include an engine and a generator, each housed within an enclosure. In operation 10904, a first electrical connector of the REEP is connected to a second electrical connector of the aircraft. In operation 10906, power is output from the generator of the REEP to at least one electrical component or electrical bus of the aircraft. In various embodiments, the REEP can also have a third set of electrical connectors such that the REEP can be electrically connected to another power consuming or distributing device that is not the aircraft. Thus, in 10908, the additional connector or the third set of connectors can be connected to another device separate from the aircraft to provide power to the other device without disconnecting the REEP from the aircraft. In various embodiments, the REEP can output AC power from the generator of the REEP or the REEP can include an inverter such that AC power from the generator is converted to DC power and output by the REEP through one of the connectors. In various embodiments, the REEP can be configured to output AC power through a first set of connectors and DC power through a second set of connectors.

[0037] At 10910, the REEP may be controlled based on control signals from the aircraft transmitted through the first and second electrical connectors, which may include connectors for control wiring. For example, one or more control wires / connectors may be used to receive throttle control signals or power request signals from the aircraft controller at the removable energy source controller through the connectors. In this manner, the aircraft may be able to control the amount of power generated and output to the aircraft. In various embodiments, other types of control signals and / or wiring / connectors may be used between the REEP and the aircraft. For example, status signals from the REEP may be transmitted to the aircraft, such as different sensor readings within the REEP (e.g., temperature, fuel level, current power output, etc.). Other signals, such as on / off signals to turn the REEP on or off, whether it outputs AC or DC power and how much, respectively, may be transmitted from the aircraft to the REEP to control the REEP.

[0038] At 10912, the REEP is powered down such that power is no longer being output from the REEP to the aircraft. At 10914, the REEP's electrical connector is disconnected from the aircraft's electrical connector. At 10916, the REEP is removed from the aircraft. In this manner, method 10900 demonstrates how a REEP can be attached to an aircraft to provide power to the aircraft, and how the REEP can be made removable from the aircraft. In this manner, the REEP can be a temporary, removable power source for the aircraft.

[0039] As such, various embodiments of a REEP are described herein that may be attached to an aircraft (using either non-destructively removable or destructively removable components) to provide electrical energy used to drive propulsion systems elsewhere on the aircraft (e.g., propulsion motors, rotors, etc. that are not part of the REEP or within the REEP housing / enclosure). Thus, such an aircraft may function with the REEP removed (e.g., the aircraft may have its own built-in or internal energy source, but the REEP provides an external, additional source of energy to power the aircraft). The REEP may be connected to other components, such as, for example, an electric bus and / or wiring within the aircraft, where at least one electric propulsion motor and / or at least one propulsion battery (e.g., batteries used to power the propulsion motors) are attached to the electric bus and / or other components. In this manner, the REEP may be configured to augment electrical energy already contained in other portions of the aircraft. As such, the REEP may provide energy to the same propulsion motors, which may also be powered by batteries or other energy sources on the aircraft other than where the REEP is mounted or attached. Thus, the REEP can function as an external energy source that supplements or augments power already available to the aircraft from its own energy sources or internal energy sources. Thus, the REEP or external energy source may enable the aircraft to fly or be flyable with or without the REEP / external energy source attached to the aircraft via a structural connection.

[0040] Flexible Architectural Elements Aircraft typically have custom-designed propulsion mechanisms and methods for powering those propulsion mechanisms. In this way, the propulsion mechanisms and the power provided to those propulsion mechanisms can be optimized to provide the amount of propulsion required for a particular type and size of aircraft while minimizing the weight of components within the aircraft. In other words, the propulsion mechanisms and the power of those propulsion mechanisms are often optimized for a particular type and size of aircraft, so that components of one aircraft cannot easily be used with different types of aircraft drive architectures, such as direct drive aircraft, parallel drive aircraft, and serial drive aircraft.

[0041] Various embodiments of flexible architectures and optimized components thereof for aerospace hybrid systems are described herein. A hybrid system may be or include a system in which fuel is combusted in a piston, rotary, turbine, or other engine and the output of the piston engine is operatively connected to a generator to output electrical power. The embodiments described herein may include a flexible system capable of powering many different types of aircraft and propulsion mechanisms. Such a system may advantageously reduce the complexity of designing different types of aircraft, reduce the manufacturing costs of such systems as less customization allows for economies of scale in mass production of the systems, and ultimately reduce the complexity of aircraft using the systems described herein.

[0042] The flexible architecture described herein may further be used to power propulsion mechanisms in different ways within the same aircraft or different aircraft. For example, a flexible architecture for powering propulsion mechanisms may be capable of operating in multiple different modes to power different types of propulsion mechanisms. A first aircraft may utilize one, some, or all of the multiple different modes in which the flexible architecture can operate. A second aircraft may utilize one, some, or all of the multiple different modes, and the mode utilized by the second aircraft may be different from the mode utilized by the first aircraft.

[0043] Thus, different aircraft may utilize different modes of powering the propulsion mechanisms provided by the flexible architecture described herein. Although the use of the flexible architecture may be customized in this manner, the physical hardware of the flexible architecture may be adapted for use by different aircraft with minimal or no changes to the physical components of the flexible architecture described herein. Instead, the use of different modes in different aircraft may be achieved primarily based on how the components of the flexible architecture are controlled using a processor or controller. Thus, computer readable instructions may be stored in a memory operatively coupled to a processor or controller, such that when the instructions are executed by the processor or controller, a computing device including the processor or controller may control the various components of the flexible architecture described herein to utilize any possible modes of use desired for a particular implementation, aircraft, flight phase, etc.

[0044] Power generation and propulsion systems for aircraft may also utilize various cooling systems to ensure that various components of the aircraft are kept at safe temperatures for operation, as well as to ensure that the components remain within temperature ranges that they can operate more efficiently. Further, described herein are advantageous cooling systems that utilize various aspects of the hybrid architecture described herein to efficiently cool components of a flexible architecture for powering the propulsion mechanisms of an aircraft.

[0045] An aircraft having hardware for providing different modes of power to the propulsion mechanisms may have various components for which it is desirable to provide cooling. Thus, a single cooling system that efficiently moves air to the various components that enable the different power modes may reduce the weight of the aircraft as well as the power consumption of the cooling system. Figures 1-8 and the following accompanying description relate specifically to an example flexible architecture for powering the propulsion system of an aircraft. Figures 9-21 and the following accompanying description relate to various embodiments of a cooling system for the example flexible architecture.

[0046] 9A illustrates an example of an aerospace hybrid system flexible architecture 101 in accordance with an illustrative embodiment. As discussed herein, the flexible architecture 101 can be efficiently used in a wide range of applications with a single hybrid generator system that can be adapted in multiple ways (e.g., used in different modes) depending on the aircraft requirements and flight phase.

[0047] The flexible architecture 101 of FIG. 9A is a hybrid generator including an engine 105, a clutch 115, a generator / motor 121, and a power shaft 111. As described further below, the flexible architecture 101 can be used to achieve a variety of different modes as needed depending on the installation requirements of a particular aircraft or a particular flight phase. The engine 105 can be a combustion engine, such as an internal combustion engine. More specifically, the engine 105 can be one of a piston internal combustion engine, a rotary engine, or a turbine engine. Such engines can use standard gasoline, jet fuel (e.g., Jet A, Jet A-1, Jet B fuel), diesel fuel, biofuel alternatives, and the like. In various embodiments, other types of engines can be used, such as small engines (e.g., Rotax gasoline engines) in drone implementations.

[0048] As mentioned above, the engine 105 may be a piston combustion engine. A piston combustion engine may advantageously rotate an output rotor or shaft at revolutions per minute (RPM) that is more desirable for direct output to power a generator and / or a propulsion mechanism (e.g., a propeller) than other engines. For example, a piston combustion engine may have an output on the order of several thousand RPM. For example, the output of a piston combustion engine may be in the range of 2200-2500 RPM, which may be a desirable RPM for a propeller. In particular, the propeller may be designed to have a size that results in a desired tip speed of the propeller based on the RPM output of the piston combustion engine (e.g., 2200-2500 RPM). Other types of engines, such as turbine engines, may output rotational power on the order of tens of thousands of RPM, much higher than a piston combustion engine. Alternative embodiments may drive a motor / generator at the higher RPM of the turbine engine to benefit efficiency, power output, or other important factors. In some embodiments, a gearbox can be added between the output of the high RPM engine and the other components of FIG. 9A to reduce the output RPM of the engine 105. However, adding a gearbox may also increase the weight of the system, which may be undesirable in some embodiments. Piston combustion engines may have an additional advantage in terms of noise compared to turbine engines. Turbine engines are typically louder than piston combustion engines, and the human-perceived noise from turbine engines is typically more annoying to the listener than the noise generated by piston combustion engines. In urban or densely populated environments where reduced noise is desired, a quieter engine may be more valuable.

[0049] The engine 105 can output rotational power to the clutch 115, which can be controlled to engage or disengage the power shaft 111. In other words, the power shaft 111 can be engaged with the rotational output of the engine 105 by the clutch 115, whereby rotational power can be transferred between the output of the engine 105 and the power shaft 111. When the clutch 115 is disengaged from the output of the engine 105 and the power shaft 111, the power shaft 111 can rotate independently of the output of the engine 105. The clutch 115 can be physically located between the engine 105 and the generator / motor 121, and can even contact the opposite side of the engine 105 and the generator / motor 121 to reduce the overall footprint of the flexible architecture. In FIG. 1A, the clutch 115 is further described herein and shown in other figures. However, in various embodiments, any mechanism capable of releasably decoupling the engine 105 and the power shaft 111 can be used in addition to or in place of the clutch. For example, the decoupling can be based on absolute revolutions per minute (RPM) or relative RPM between the output of the engine 105 and the power shaft 111, such as in an overrunning clutch.

[0050] The generator / motor 121 may also be engaged with the power shaft 111 or disengaged from the power shaft 111. In other words, the generator / motor 121 may be controlled to be switched off so that the rotation of the power shaft 111 does not cause the generator / motor 121 to generate power. Similarly, the generator / motor 121 may be controlled to be switched on so that the rotation of the power shaft causes the generator / motor 121 to generate power. The generator / motor 121 is referred to as a generator / motor because it can function as both a generator and a motor. In various embodiments, the generator / motor 121 may be referred to as an electric machine, which may be a generator, an electric motor, or both.

[0051] The flexible architecture further includes a power input / output (I / O) 125 connected to the generator / motor 121. As described further herein, the generator / motor 121 may generate power based on the rotation of the power shaft 111 that is output via the power I / O 125, or may receive power via the power I / O 125 that may be used to drive the power shaft 111. The wiring of the power I / O 125 may include multiple wires. In various embodiments, the wiring for inputting power to the generator / motor 121 may be the same as the wiring used to output power from the generator / motor 121. In various other embodiments, a first wiring may be used for inputting power and a different second wiring may be used for outputting power (using different wiring for input and output). In various embodiments, the generator / motor 121 may also have wiring that is used to control the generator / motor 121 and is connected to relay sensor or other data regarding the operation of the generator / motor 121 to a controller or the like.

[0052] The generator / motor 121 can also function as a driver for the power shaft 111. Upon receiving power via the power I / O 125 from a battery or some other form of electrical energy storage elsewhere in the system, the generator / motor 121 can provide a rotational force to drive the power shaft 111. This can occur as long as the generator / motor 121 is controlled to be switched on so that it is engaged with the power shaft 111. If the generator / motor 121 is controlled to be switched off so that it is not engaged with the power shaft 111, the power shaft 111 may no longer be rotated by the generator / motor 121.

[0053] The power output from the power I / O 125 may be used to drive electric motors for electric propulsion (e.g., propellers). The power output from the power I / O 125 may also be used to power and / or charge other devices on the aircraft or aerospace vehicle. For example, the power output from the power I / O 125 may be used to charge one or more batteries. The power output from the power I / O 125 may also be used to power other devices or accessories on the aircraft or aerospace vehicle. Because the power I / O 125 also has an input, the power shaft 111 may be driven by any power received via the power I / O 125, such as power from one or more batteries. The power generated by the generator / motor 121 may be alternating current (AC) power. The AC power may be converted to direct current (DC) power by power electronics (e.g., a rectifier or inverter) and output to a DC bus. This DC bus may be connected to a battery and / or an electric propulsion. In this manner, the electric propulsion may be powered via the DC bus. In various embodiments, the motors of the electric propulsion may use AC power, and thus the DC power from the DC bus may be converted from DC power to AC power before being used by the electric propulsion (e.g., an inverter). In various embodiments, the AC power generated by the generator may be supplied directly to the motors or other devices without being converted back to DC power again. In such embodiments, such AC power may be transmitted over an AC power bus or similar wiring.

[0054] Any rotation of the power shaft 111 itself may also be used to drive one or more propulsion mechanisms, whether driven by the engine 105 or the generator / motor 121. For example, the rotation of the power shaft 111 may be used to directly drive a propeller or may be used to power an electric motor that drives the propulsion mechanism. The rotation of the power shaft 111 may also drive a gearbox that is operatively connected to another component, such as one or more propellers, one or more rotors, or other rotating devices for various applications on the aircraft.

[0055] The accessory pad 130 may also be coupled to the engine 105 and may include a low voltage direct current (DC) generator for power separate from the generator / motor 121 and power I / O 125, which may be configured for high voltage and high power I / O. In some embodiments, the generator / motor 121 may have two different windings and the power I / O 125 may have two different outputs (e.g., high voltage and low voltage). An accessory power source may be associated with one of the outputs of the power I / O 125 in addition to or instead of the output of the accessory pad 130. The accessory pad 130 may be used to power devices or accessories on the aircraft or aerospace vehicle that do not require the high voltage or current output that may be output by the generator / motor 121 at the power I / O 125. The voltage (HV) of the aircraft may be, for example, 400 volts (V) or 800V, but could be anywhere between 50V and 1200V. The low voltage (LV) of the aircraft may be 12V, 14V, 28V, or other voltage less than 50V.

[0056] Figure 9B illustrates an additional example of a flexible architecture 150 for an aerospace hybrid system in accordance with an illustrative embodiment. In particular, the flexible architecture 150 of Figure 9B includes several components that may be the same or similar to those described above with respect to Figure 9A, where the flexible architecture 150 includes an engine 155, a clutch 175, a power shaft 180, and / or a generator / motor 185. The flexible architecture 150 further illustrates an output of the engine 155 in the form of a crankshaft 160, which is rigidly connected to an output flange 165. The output flange 165 is rigidly connected to one side of the clutch 175 using bolts 170.

[0057] The clutch 175 may be configured to engage the power shaft 180 to convert rotational motion from the crankshaft 160 and the output flange 165 to the power shaft 180. The clutch 175 may further be configured to disengage from the power shaft 180, thereby allowing the power shaft 180 to rotate independently relative to the crankshaft 160 and the output flange 165. Additionally, FIG. 9B illustrates how all of the rotatable components of the flexible architecture 150 may be aligned along a single axis 190. The rotatable components of FIG. 9A may similarly be aligned along a single axis, as shown in FIG. 9B. Additionally, the power shaft 180 may be a splined shaft that fits into the inner diameter openings of the clutch 175 and the generator / motor 185. Other features other than splines, such as tapers, may also be used. In either case, the generator / motor 185 and / or the clutch 175 may be configured to match and connect to the splines, tapers, or other features on the power shaft 180 so that the components can properly engage with one another.

[0058] In various embodiments, the clutch 175 may be a different type of clutch or other mechanism capable of decoupling the power shaft 180 from the output of the engine 155. For example, the clutch 175 may be a plate-type clutch and may be a dry or wet clutch. Such a plate-type clutch may be engaged / disengaged or otherwise controlled mechanically, hydraulically, and / or electrically (e.g., by the controllers 205, 220, and / or 280 of FIGS. 10A and 10B). A plate-type clutch may have a variety of numbers of plates, such as three, five, or ten plates. In various embodiments, the clutch 175 or other clutches described herein may be a one-way clutch, an overrunning clutch, or a sprag clutch. A one-way clutch or a sprag clutch may be configured to disengage the engine output from the power shaft while the electric machine rotates the power shaft faster than the engine output. In other words, when the engine 155 is outputting less power to the power shaft 180 than the generator / motor 185, the clutch 175 can automatically mechanically disengage the output of the engine 155 from the power shaft 180, e.g., without any electrical control input being used to accomplish that disengagement. When the engine 155 has a higher RPM or outputs more power than the generator / motor 185, a one-way or sprag clutch is engaged so that power is added from the output of the engine 155 to the power shaft 180. Another type of clutch that can be used is a centrifugal clutch, where as the RPM increases, the weight of the plates of the clutch gradually activates one or more levers, compressing the plates of the centrifugal clutch and engaging the plates to connect, e.g., the output of the engine 155 and the power shaft 180.

[0059] Advantageously, the generator / motor 121 of FIG. 9A and / or the generator / motor 185 of FIG. 9B may be used as a starter for the engine 105 or the engine 155, respectively. In other words, the generator / motor 185 may be used to rotate the crankshaft 160 while the clutch 175 is engaged to start the engine 155. Such a system may be advantageous, for example, where the generator / motor 185 may be powered by a battery or other power source. Thus, the engine 155 may be a piston combustion engine as described herein, and may not require a separate starter component, reducing the weight and complexity of the flexible architecture described herein.

[0060] FIG. 10A illustrates a block diagram depicting an aircraft control system 200 used with a flexible architecture 201 for an aerospace hybrid system, according to an exemplary embodiment. The aircraft control system 200 may be used to realize one or more of the various modes discussed below, for example, in which the flexible architecture described herein may be used. The flexible architecture 201 may have the same, similar, or some or all of the components of the flexible architectures 101 and / or 150 of FIG. 9A and / or FIG. 9B. The aircraft control system 200 may include one or more processors or controllers 205 (hereinafter referred to as controller 205), memory 210, an aircraft main controller 220, an engine 230, a generator / motor 235, a clutch 240, a power I / O 245, an accessory pad 250, and one or more sensors 260. The connections in FIG. 10A illustrate control signal related connections between the components of the aircraft control system 200. Other connections not shown in FIG. 10A may exist between different aspects of the aircraft and / or aircraft control system 200 to provide power, such as high voltage (HV) or low voltage (LV) power for the aircraft.

[0061] The memory 210 may be a computer-readable medium configured to store instructions. Such instructions may be computer-executable code executed by the controller 205 to implement various methods and systems described herein, including various modes and combinations of modes using the flexible architecture herein. The computer code may be written such that various methods of implementing different modes of the flexible architecture herein are automatically implemented based on various inputs, for example, indicative of a particular flight phase (e.g., landing, takeoff, cruise, etc.). In various embodiments, the computer code may be written to implement various modes herein based on inputs from a user or pilot of the aircraft or aerospace vehicle, or may be implemented based on a combination of user inputs and automatic implementations based on non-human inputs (e.g., from sensors on or off the aircraft, based on a planned flight plan, etc.). The controller 205 may be powered by a power source on the aircraft or aerospace vehicle, such as the accessory pad 130, one or more batteries, an output of the power I / O 125, a power bus of the aircraft powered by any power source, and / or any other available power source.

[0062] The controller 205 may also communicate with each of the engine 230, the generator / motor 235, the clutch 240, the power I / O 245, the accessory pads 250, and / or the sensors 260. In this manner, the components of the flexible architecture may be controlled to achieve the various modes described herein. In various embodiments, the engine 230, the generator / motor 235, the clutch 240, the power I / O 245, and the accessory pads 250 may be similar to or similarly named components shown in and described above with respect to FIG. 9A. The power I / O 245 may also include pre-charge electronics to protect the electrical components of the flexible architecture, including, for example, a direct current (DC) bus, from excessive inrush current during startup as described herein. For example, if a high voltage (HV) bus is 400V and a new component is connected to the 0V HV bus, the instantaneous inrush current may be very large and may cause damage to the HV bus and / or the component. As a result, the precharge electronics can slowly increase the voltage of the component before fully connecting it to the HV bus or other power source.

[0063] The sensors 260 may include various sensors for monitoring different components of the flexible architecture 201. Such sensors may include, for example, temperature sensors, tachometers, fluid pressure sensors, voltage sensors, current sensors, status sensors, etc., to determine the current state of the clutch 240, or other types of sensors. For example, voltage and / or current sensors may be used to indicate the function and settings of the motor / generator, to inform the clutch of the selected state, or to adjust other components of the system. Status sensors may also indicate the particular mode in which the flexible architecture is being used, and the system may receive input (e.g., from the pilot, from an automatic flight control) to change the system to a different state or mode for a particular upcoming flight phase. Other sensors may include a pitot tube to measure the airspeed of the aircraft, an altimeter to measure the altitude of the aircraft, and / or a global positioning system (GPS) or similar geographic location sensor to determine a position relative to the ground and / or known / mapped structures.

[0064] Components within the dashed lines of flexible architecture 201 in FIG. 10A may be associated with the flexible architecture described herein, while aircraft main controller 220 may be associated with a broader aircraft system. In other words, aircraft main controller 220 may control aspects of the aircraft other than flexible architecture 201, while controller 205 may control aspects of the aircraft related to flexible architecture 201. Aircraft main controller 220 and controller 205 may communicate with each other to coordinate the powering of various propulsion mechanisms of the aircraft. For example, aircraft main controller 220 may send signals to control device 205 requesting a particular power output level for one or more particular propulsion mechanisms. Controller 205 may receive such control signals and determine how to coordinate flexible architecture 201 (e.g., which mode to enter, how to control elements of flexible architecture 201) to output the desired power level based on the control signals from aircraft main controller 220. In various embodiments, aircraft main controller 220 may send signals related to the control of particular aspects of flexible architecture 201. In other words, in addition to or instead of transmitting a desired power output signal to controller 205, controller 205 can function as a repeater to retransmit control signals from aircraft main controller 220 to the components of flexible architecture 201, and controller 205 determines from the control signals how to control the individual components of flexible architecture 201.

[0065] In various embodiments, aircraft main controller 220 may also transmit control signals related to future desired power output, future flight phases, flight plan information, etc. In this manner, controller 205 may receive information regarding the anticipated power requirements of the aircraft and use it to determine how to control aspects of flexible architecture 201, both now and in the future. For example, flight plan information may be used to determine when to use battery power, when to charge batteries, etc. In another example, if a large power demand is anticipated, controller 205 may ensure that engines 230 are running before beginning to provide the desired level of power.

[0066] In various embodiments, controller 205 may also communicate with one or more batteries to monitor their charge levels, control when the batteries are charged or discharged, control when the batteries are used to power generators / motors 235, and control when the batteries are used to directly power other aspects of the aircraft. However, in other embodiments, aircraft main controller 220 may communicate with the aircraft's batteries and / or relay information regarding the batteries and their control to controller 205. Similarly, in cases where the aircraft's batteries are controlled using aircraft main controller 220 rather than controller 205, controller 205 may send control signals related to the batteries to the aircraft main controller such that the batteries may be controlled as necessary or desired for the functioning of flexible architecture 201.

[0067] In various embodiments, the power I / O 245 may include two different outputs (e.g., a high voltage (HV) output and a low voltage (LV) output) associated with two different windings of the generator / motor 235. Thus, two different voltages (e.g., HV and LV) may be output and controlled by the controller 205 and / or the aircraft main controller 220. The power I / O 245 may additionally or alternatively have voltage conversion components (e.g., DC / DC converters) to enable it to output two or more different voltages. In such an embodiment, the two different outputs may be achieved without using two separate windings. The two different outputs may be output to different power buses of the aircraft, such as, for example, an HV bus and an LV bus. The two outputs of the power I / O 245 may also be independently controlled by the controller 205. Thus, the outputs may be turned off (e.g., by turning off the field current of the motor / generator, thereby allowing the generator's power shaft and rotor to rotate or freewheel relative to the rest of the motor / generator).

[0068] In some embodiments, the accessory pads may not be controlled by controller 205 and / or aircraft main controller 220. The accessory pads may simply be on all the time when engines 230 are running, or may be controlled separately (e.g., by a manual switch toggled by a user) to control when and how power is provided to the aircraft accessories.

[0069] In some embodiments, controller 205 may communicate with a wireless transceiver that may be onboard an aircraft or aerospace vehicle, thereby allowing controller 205 to communicate with other computing devices that are not hardwired to system 200. In this manner, instructions or inputs for implementing the various modes of the flexible architecture described herein may also be received wirelessly from remote computing devices. In other embodiments, system 200 may communicate only with components onboard an aircraft.

[0070] Figure 10B illustrates a block diagram depicting a second aircraft control system 275 used with the flexible architecture for an aerospace hybrid system, according to an example embodiment. In the example of Figure 10B, system 275 does not have a separate aircraft main controller as in Figure 10A. Instead, the entire aircraft has the flexible architecture and a single main controller 280 that controls all aspects of the aircraft (including, for example, the aircraft's propulsion mechanism 255).

[0071] The controller 285 can communicate with and control one or more propulsion mechanisms 255 on the aircraft. The controller 285 can also communicate with one or more sensors 270 on the aircraft or aerospace vehicle, which may be aircraft sensors and flexible architecture sensors. In particular, the sensors 260 can be embedded in any of the components of Figures 9A and / or 9B described above, and thus can be used to inform how the devices of Figures 9A and / or 9B are controlled and / or how the modes described herein are implemented as described herein.

[0072] 10A or 10B, controller 205, controller 285, and / or aircraft main controller 220 may also be in communication with a cooling system configured to cool and / or heat any component of the flexible architecture, one or more batteries, or other aspects of the aircraft, such that the cooling system may also be controlled in conjunction with other systems and methods described herein.

[0073] Five particular modes that may be realized using various embodiments of the flexible architectures described herein (including, for example, the flexible architectures shown in and described with respect to Figures 9A, 9B, 10A, and 10B) are described below.

[0074] In a first mode, which may be referred to herein as a hybrid generator mode, a clutch (e.g., clutch 115 of FIG. 9A and / or clutch 175 of FIG. 9B) may be controlled to engage an engine (e.g., engine 105 of FIG. 9A and / or engine 155 of FIG. 9B) with a power shaft (e.g., power shaft 111 of FIG. 9A and / or clutch output / power shaft 180) extending from the clutch to a generator / motor (e.g., generator / motor 121 of FIG. 9A and / or generator / motor 185 of FIG. 9B), whereby the engine rotates a power shaft in the generator / motor to generate electrical power that is supplied to other systems on the aircraft, such as propulsion mechanisms / systems, via a power I / O (e.g., power I / O 125 of FIG. 9A). For example, such propulsion mechanisms / systems may be powered using an electric motor, and the electrical power output by the generator / motor in the first mode may be used to drive such propulsion mechanisms / systems. That is, in a first mode, the clutch can be used to engage the engine with a power shaft to drive a generator / motor, which can then output electrical power.

[0075] In a second mode, which may be referred to herein as a direct drive engine mode, a clutch (e.g., clutch 115 of FIG. 9A and / or clutch 175 of FIG. 9B) may engage an engine (e.g., engine 105 of FIG. 9A and / or engine 155 of FIG. 9B) output to a power shaft (e.g., power shaft 111 of FIG. 9A and / or clutch output / power shaft 180) that extends through a generator / motor (e.g., generator / motor 121 of FIG. 9A and / or generator / motor 185 of FIG. 9B) to provide mechanical power to a propulsion mechanism, such as a propeller of an aircraft. In such a mode, the magnetic field may be removed from the generator / motor (e.g., the generator / motor may be controlled to be off or disengaged) such that the generator / motor power shaft and rotor rotate or freewheel and thus the generator / motor power I / O (e.g., power I / O 125 of FIG. 9A) is disengaged and does not output power. That is, in the second mode, the engine can drive the power shaft to mechanically or otherwise power the propulsion mechanism, while the power shaft rotates in the generator / motor without receiving or outputting power at the power I / O.

[0076] In a third mode, which may be referred to herein as an augmented thrust mode, a clutch (e.g., clutch 115 of FIG. 9A and / or clutch 175 of FIG. 9B) may engage an engine (e.g., engine 105 of FIG. 9A and / or engine 155 of FIG. 9B) to a power shaft (e.g., power shaft 111 of FIG. 9A and / or clutch output / power shaft 180) that extends through a generator / motor (e.g., generator / motor 121 of FIG. 9A and / or generator / motor 185 of FIG. 9B), which is used as a motor drawing power via a power I / O (e.g., power I / O 125 of FIG. 9A) from an external source such as a battery pack. This provides a higher mechanical power output to the power shaft than the engine or generator / motor can provide. That is, in the third mode, both the engine and the generator / motor are used to simultaneously drive the power shaft to deliver power (electricity) to the propulsion mechanism.

[0077] In a fourth mode, which may be referred to herein as a direct drive generator / motor mode, a clutch (e.g., clutch 115 of FIG. 9A and / or clutch 175 of FIG. 9B) may disengage the engine (e.g., engine 105 of FIG. 9A and / or engine 155 of FIG. 9B) from the generator / motor (e.g., generator / motor 121 of FIG. 9A and / or generator / motor 185 of FIG. 9B) such that power may be provided to the generator / motor via the power I / O (e.g., power I / O 125 of FIG. 9A) to drive the generator / motor as a motor and provide mechanical power to the power shaft (e.g., power shaft 111 of FIG. 9A and / or clutch output / power shaft 180). That is, in the fourth mode, only the generator / motor may provide power (electricity) to the propulsion mechanism based on the power received at the power I / O.

[0078] In a fifth mode, which may be referred to herein as an engine power split mode, a clutch (e.g., clutch 115 of FIG. 9A and / or clutch 175 of FIG. 9B) may engage an engine (e.g., engine 105 of FIG. 9A and / or engine 155 of FIG. 9B) with a generator / motor (e.g., generator / motor 121 of FIG. 9A and / or generator / motor 185 of FIG. 9B) such that the engine can not only rotate the generator / motor as a generator to provide power to other systems on the aircraft via the power I / O (e.g., power I / O 125 of FIG. 9A), but also provide mechanical power to a power shaft (e.g., power shaft 111 of FIG. 9A and / or clutch output / power shaft 180) to drive a system such as a propeller. That is, in the fifth mode, the engine may be used to drive a power shaft and generator / motor to output power (electricity) via the power I / O and power shaft.

[0079] As described herein, any of these five modes (or variations thereof) may be used with the single flexible architecture described herein. Furthermore, particular modes and / or combinations of modes may be beneficial for particular aircraft or aerospace vehicle types, particular propulsion types, particular flight phases of an aircraft or aerospace vehicle, etc.

[0080] For example, in a hybrid electric vertical take-off and landing (VTOL) aircraft with electric motor-driven propellers, the flexible architecture herein may be used only as a power source such that the flexible architecture can power the aircraft in a first mode (e.g., hybrid generator mode) during any portion of the flight phase in which the aircraft's power bus or one or more of the aircraft's motors must be powered.

[0081] In another example, in an aircraft with a single large main pusher propeller (e.g., at the rear of the aircraft fuselage) and an array of electric motors / propellers (e.g., on the aircraft wings), the flexible architecture can be used in a fifth mode (e.g., engine power split mode) during takeoff to mechanically power the main pusher propeller and electrically power the wing-mounted motors. Figures 11 and 12 show two examples of such aircraft 300 and 400 that can use the flexible architecture for aerospace hybrid systems according to exemplary embodiments. For example, aircraft 300 has a main pusher propeller 305 and aircraft 400 has a main pusher propeller 405 in the form of a ducted propulsion fan. In both examples, the fifth mode described herein can be used to mechanically power the main pusher propellers 305 and 405 from a power shaft. Additionally, the wing-mounted electric motors / propellers 310 and 410 can be driven with electrical power from a motor / generator as described herein.

[0082] Alternatively, the flexible architecture described herein may be used to power configurations such as those shown in FIGS. 11 and 12 in a third mode (e.g., augmented mode) by the battery pack powering both wing-mounted motors at takeoff to augment engine power to the power shaft driving the main propulsion propeller. During cruise flight, the aircraft may use the second mode (e.g., direct drive engine mode) to drive only the main propulsion propeller. In another example, during cruise flight, the aircraft may include a clutch between the power shaft and the propulsion propeller, and the controller may operate the aircraft in a first mode (e.g., hybrid generator mode) driving the wing-mounted motor by disengaging the power shaft from the propulsion propeller and outputting power from the generator / motor to the wing-mounted motor. In another example (e.g., emergency situations such as engine failure), the propulsion prop may be driven in a fourth mode (e.g., direct drive generator / motor mode) using power input to a power I / O such as one or more batteries.

[0083] In another example, the aircraft may be a VTOL aircraft with a gyrocopter-type main rotor that can operate powered or unpowered and may have wing-mounted forward thrust motors and propellers. In one embodiment, the flexible architecture may be used entirely in a first mode (e.g., hybrid generator mode) where power provided from the power input / output (and generator / motor) drives motors coupled to the gyrocopter-type main rotor and power is used to drive wing-mounted motors. In one embodiment, the aircraft may be configured with a clutch between the power shaft and the gyrocopter-type main rotor such that the flexible architecture may use a second mode (e.g., direct drive engine mode) or a third mode (e.g., augmented thrust mode) to spin the gyrocopter-type main rotor (e.g., to bring the gyrocopter-type rotor up to takeoff speed). In such an example, the controller may switch the flexible architecture to the first mode (e.g., hybrid generator mode) after the gyrocopter-type rotor has reached speed (e.g., switch to the first mode for cruise flight). A fourth mode (e.g., direct drive generator / motor mode) can be used again in the event of engine failure, using electrical power to drive a powered shaft (and thus a gyrocopter-type rotor) from a power source such as one or more batteries.

[0084] 13 illustrates another example aircraft 500 that may employ a flexible architecture for an aerospace hybrid system, according to an example embodiment. For example, the aircraft 500 may include multiple (e.g., eight) electric motor / propellers 505 on a tilt wing that may be powered using a first mode (e.g., hybrid generator mode) described herein, in which an engine may be engaged with a power shaft using a clutch to drive a generator / motor that may output power to the various electric motor / propellers 505 on the tilt wing.

[0085] Thus, described herein is an advantageous flexible architecture for an aircraft that can achieve various modes for powering the propulsion. While a particular aircraft and propulsion configuration may not utilize each of the modes described herein for which the flexible architecture is capable, the flexible architecture can still be implemented in different aircraft to achieve different modes. Similarly, while one example of a flexible architecture having five different modes for powering the propulsion is described in detail herein, other flexible architectures having fewer, more, or different modes for powering the propulsion are also contemplated herein.

[0086] For example, the flexible architecture may not have a clutch as described herein and still achieve the various modes described herein when it is desirable to couple the engine output to the motor / generator and / or power output shaft of the system. For example, in a first mode, the engine may rotate a power shaft to generate electricity via a generator. In a second mode, the engine may directly drive, for example, mechanical propulsion components, but it is not necessary to disengage the engine from the motor / generator or power shaft because the motor / generator can be turned off or the motor / generator power shaft and rotor can be allowed to freewheel within the motor / generator. In a third mode, the engine and motor / generator are used to drive the power shaft, so it is not desirable to disengage the engine and motor / generator using a clutch. In a fifth mode, the engine may rotate a power shaft to generate electricity via a generator, which mechanically powers the propulsion mechanism. Thus, in an aircraft utilizing any of the first, second, third, and / or fifth modes described above, it is not necessary to disengage the power shaft from the engine output. Thus, in embodiments using any combination of the first, second, third, and / or fifth modes (but not the fourth mode), a clutch may not be used since the system may always connect the engine output to the motor / generator power shaft, which may be of value since clutches may be heavy and / or unreliable.

[0087] 14 is a flow chart illustrating a first exemplary method 601 for using a flexible architecture of an aerospace hybrid system in different flight phases of an aircraft with a main propulsion propeller, according to an exemplary embodiment. In particular, the aircraft may be an aircraft with a single larger propulsion propeller and an array of electric motors and corresponding smaller propellers on the wings. During the takeoff flight phase at 603, the fifth mode described herein may be used to mechanically power the main propulsion propeller and to provide electrical power to the wing-mounted motors. During the cruise flight phase at 605, the second mode described herein may be used to mechanically power only the main propulsion propeller and not provide electrical power to the smaller electric motors / propellers.

[0088] FIG. 15 is a flow chart illustrating a second exemplary method 700 for using a flexible architecture of an aerospace hybrid system in different flight phases of an aircraft with a main propulsion propeller, according to an exemplary embodiment. In particular, the aircraft may be an aircraft with a single larger propulsion propeller and an array of electric motors and corresponding smaller propellers on the wings. During the takeoff flight phase at 702, a third mode described herein, called thrust augmentation, may be used to power the main propulsion propeller via a generator / motor (drawing power from the battery) and to mechanically power the main propulsion propeller directly from the engine. Additionally, power (generated by the generator / motor and / or generated directly from the battery) may also be provided to the electric motors on the wings during takeoff. During the cruise flight phase at 704, a second mode described herein may be used to mechanically power only the main propulsion propeller and not the smaller electric motors / propellers.

[0089] 9A , when engine 105 applies power to power shaft 111 and clutch 115 is engaged such that generator / motor 121 is not running or turned on, power shaft 111 may freewheel within generator / motor 121 (e.g., in the second mode described above). Similarly, power shaft 180 of FIG. 9B may freewheel within generator / motor 185 in various embodiments. However, engine 105 and / or engine 155 may generate torque pulses on power shaft 111 and / or power shaft 180 that may be dangerous to generators such as generator / motor 121 and / or generator / motor 185 when clutch 115 and / or clutch 175 engage the respective power shafts 111 and / or 180. In other words, a large torque pulse on the shaft, similar to that which can occur when certain types of engines (e.g., diesel piston internal combustion engines) fire, can cause high angular accelerations that can cause fatigue or damage to the components of the generator / motor 121 and / or generator / motor 185 coupled to the power shafts 111 and / or 180. As such, components to mitigate this torque can smooth the torque on the power shafts 111 and / or 180, such as using a flywheel or other heavy damping or spring coupling system.

[0090] FIG. 16A illustrates an example of a flexible architecture 800 for an aerospace hybrid system having a flywheel for absorbing vibration torque, according to an exemplary embodiment. In particular, the flexible architecture 800 includes similar or identical components as those illustrated in and described with respect to FIG. 9B, but includes a flywheel 195 rigidly connected to an output flange 165 using bolts 170. The flywheel 195 is further rigidly connected to one side of a clutch 175 by bolts 198. Thus, rotational motion can be transferred from the engine 155 through the crankshaft 160, the output flange 165, and the flywheel 195 to the clutch 175. The clutch 175 can engage or disengage the power shaft 180 to selectively convert the rotational motion received from the flywheel 195 to the power shaft 180. The flywheel 195 can also be, for example, a dual mass flywheel or a spring coupling.

[0091] In various other embodiments, a flywheel may not be used. For example, further embodiments of damping systems and damping devices are described herein that may dampen torque on a power shaft (e.g., power shaft 111) but do not include a flywheel. Additionally, in various embodiments, a flywheel and other damping systems or components may be used in combination to dampen or smooth the torque applied to the power shaft.

[0092] For example, a power shaft or rotor within the generator / motor itself may be rigidly coupled to the crankshaft of the generator / motor. In this manner, the crankshaft and rotor may together dampen torque pulses on the power shaft or rotor, reducing tangential acceleration due to torque pulses from the engine. In such an embodiment, a clutch may be omitted. As such, the damping system is internal to the generator / motor, and the damping system's footprint and weight may be smaller than a flywheel or other damping system that may be external to the generator / motor. In particular, a rigid coupling of the power shaft or rotor to the crankshaft may increase the inertia of the power shaft or rotor, whereby that additional inertia helps to prevent the power shaft from slowing down or otherwise rotating in a manner that is susceptible to acceleration due to torque pulses of the engine. In such an embodiment, the power shaft or rotor and crankshaft may function similarly to a flywheel.

[0093] In various embodiments, a generator / motor may be used that has a stationary inner portion and a rotating outer portion, which increases the inertia of the rotating portion and may allow magnets in the generator / motor to rotate and avoid being dislodged by the Torx spike. In other words, since the magnets may already be rotating on the outer portion, a constant stabilizing radial force may be applied in addition to the tangential inertial force due to the torque spike acceleration.

[0094] The torque damping system may also be configured as part of the power shaft or rotor that connects the engine output to the generator / motor. For example, the hub between the generator / motor power shaft or rotor may include a coupling with torsional spring and / or damping properties. The torsional damping coupling may include an elastomeric part or spring (e.g., made of steel or another metal) that reduces the transmission of potentially harmful torque impulses from the engine output to the generator power shaft or rotor. The torsional damping coupling may be similar to or referred to as a resonant damping coupling. For example, such a torsional damping coupling may reduce the weight and size of the overall system as opposed to a system that uses a flywheel or other bulky damping system. One or more torsional damping couplings may be installed in the engine, between the engine and the clutch, in the clutch, between the clutch and the generator, and / or in the generator to achieve damping before the power shaft or rotor damages components of the generator itself.

[0095] Other methods of damping torque on the generator's power shaft or rotor may also be used. For example, the magnetic field on the generator may be pulsed and controlled to act on the generator's power shaft or rotor to counter some or all of the torque pulses imparted to the power shaft or rotor by the engine. Such pulses on the generator's magnetic field may be controlled based on measurements of torque pulses applied by the engine, so that generator components are not damaged by the diesel engine. For example, in the third mode described above, where both the engine and the generator / motor provide power to the power shaft, pulses may be provided from the generator to the power shaft to provide power to the power shaft and prevent damage to the generator components. In other modes described herein, the generator may be used to apply pulses to the power shaft whenever the power shaft is being driven in part by the engine overall. Thus, to adequately protect the generator's components in such a manner, the pulses applied by the generator's magnetic field to the power shaft or rotor may be configured to correlate with and appropriately counteract the torque pulses of the engine.

[0096] FIG. 16B illustrates an example of a flexible architecture 801 for an aerospace hybrid system with a flywheel and spring coupling to absorb vibration torque, according to an exemplary embodiment. In particular, the flexible architecture 801 includes similar or identical components as shown in and described with respect to FIG. 16A, but includes a spring coupling 199 rigidly connected to the flywheel 195 and the power shaft 180. The size, weight, etc. of the flywheel 195 and the characteristics of the spring coupling 199 can be adjusted according to the power output of the engine 155 and the characteristics of each other, so that the vibration torque can be reduced as desired and / or as possible. For example, different engines may generate different amounts of vibration torque, so various embodiments herein include flywheels and / or spring couplings with different characteristics to reduce vibrations transmitted from the crankshaft 160 to the power shaft 180. In various embodiments, the flexible architecture 801 may not have a clutch, such that the crankshaft 160 and the power shaft 180 are always coupled to each other. In various embodiments, a flexible architecture similar to that of FIG. 16B may also include a clutch such that the output of the engine 155 can eventually be releasably disengaged from the power shaft 180. In various embodiments, such a clutch may be connected between the spring coupler 199 and the power shaft 180, or the power shaft may be split into multiple shafts with clutches connected to the multiple shafts, or the clutch may be located anywhere between the engine 155 and the generator / motor 185 such that the output of the engine 155 can be selectively disengaged from the portion of the power shaft 180 that travels through the generator / motor 185. In various embodiments, a clutch may additionally or alternatively be located after the generator / motor 185 such that the power shaft 180 can be disengaged from the load (e.g., the propulsion system of the aircraft).

[0097] Further, examples of how the flexible architectures described herein can be packaged and / or used in actual aircraft are described below. For example, certain aircraft may use electric motors to drive their propulsion systems and therefore must have sufficient on-board electrical energy to drive those propulsion systems or be provided with a method to generate such on-board electrical energy. Furthermore, regulations in certain jurisdictions may require sufficient reserve energy to comply with aircraft operating regulations. The flexible architectures described herein can provide such electrical energy and / or reserve energy to the propulsion systems such that the systems described herein can operate in a variety of electric aircraft. For example, embodiments herein provide efficient conversion of jet fuel (or other liquid or gas fuels) to electricity such that widely available fuel sources can be used to power electric aircraft.

[0098] Figure 17 illustrates a perspective view 901 of an example flexible architecture for an aerospace hybrid system, according to an exemplary embodiment. This hybrid unit can be used as a core powerplant for a variety of aircraft types and implementations. The hybrid unit of Figure 17 is a tightly integrated powerplant that may include some, all, and / or additional elements shown and described in Figures 9A, 9B, 10A, 10B, and / or 16A / 16B.

[0099] Additionally, the hybrid unit may include an integrated cooling system 905 for cooling various aspects of the hybrid unit, heat exchangers associated with the hybrid unit, or heat sinks such as finned attachments on any aspect of the hybrid unit. The power output 910 may be or may connect to a power shaft (e.g., power shaft 110 of FIG. 9A, power shaft 180 of FIG. 9B or FIG. 16A / FIG. 16B) such that rotational power is output from the hybrid unit to a propulsion system or other aspects of the aircraft. The electrical connector 915 may be used to output electrical power (or input electrical power) as described herein. The electrical connector 915 may be, for example, an Amphenol Surlok Plus™ connector or equivalent, or any other type of suitable connector. In this manner, a main bus, such as a direct current (DC) bus of the hybrid unit, may be connected via the electrical connector 915 (e.g., power input / output 125 of FIG. 9A, power I / O 245 of FIG. 10A or FIG. 10B). These or other connectors may also facilitate connection and control of components of the hybrid unit, such as through the use of a Controller Area Network (CAN) bus, a CAN 2.0 bus, and / or an SAE J1939 bus. Such communication buses may operate at different speeds, such as 250 kilobytes per second (kbps), 500 kbps, 1000 kbps, etc. In various embodiments, the electrical connector 915 and / or other connectors may be customized for specific applications, such as different types of aircraft and the communications and power systems they use.

[0100] By virtue of the power output 910 and electrical connector 915, the hybrid unit of FIG. 17 can output mechanical power via the power output 910 and / or output electrical power via the electrical connector 915 and a DC bus within the hybrid unit (e.g., power input / output 125 of FIG. 9A , power I / O 245 of FIG. 10A or FIG. 10B ). Similarly, just as mechanical power is received via the power output 910 to generate electricity for output via the electrical connector 915, electrical power can be received via the electrical connector 915 to drive the power output 910. For example, if the aircraft has one or more batteries, additional power from the batteries can be received via the electrical connector 915 to boost the power (electrical power) applied to the power output 910, such that the power output 910 is driven by both the engines and electrical power from the aircraft's batteries, as described herein.

[0101] The hybrid unit of FIG. 17 may further include a connector 925 for connecting the engine to a fuel source. The connector 925 may be a quick fuel connection, such as an AN6 quick fuel connection. In this manner, the engine may be supplied with fuel to power the power output 910 and / or generate electricity that is output via the electrical connector 915. The hybrid unit of FIG. 17 may further include mounting hardware 921 for mounting the hybrid unit to an aircraft. Although the mounting hardware 921 is shown on the top of the hybrid unit in FIG. 17, in other embodiments, the mounting hardware may additionally or alternatively be located on either the top, bottom, side, etc. of the hybrid unit such that the hybrid unit may be mounted to the aircraft as desired.

[0102] Figure 18 illustrates a top view 1000 of the example flexible architecture of Figure 9 in accordance with an exemplary embodiment. Figure 19 illustrates a side view 1100 of the example flexible architecture of Figure 17 in accordance with an exemplary embodiment.

[0103] Thus, the hybrid units described herein can be used to power electric or hybrid electric aircraft and can provide better power than battery packs alone. For example, hybrid units such as those shown in Figures 17-19 can provide better energy density than batteries (e.g., 5-7 times better energy density). For example, the hybrid units described herein can have an equivalent energy density of 600-1200 or more watt-hours per kilogram (Wh / kg). Additionally, the hybrid units described herein advantageously have better fuel economy than other systems (e.g., 40% better fuel economy than turbine engines) and can use readily available fuels such as Jet-A, diesel, kerosene, biofuel alternatives, or other suitable or desired fuels. In other words, the hybrid units herein can include an engine, generator, inverter, and thermal management using air cooling in a compact package, thereby allowing aircraft equipped with a flexible architecture to advantageously utilize these components as power generation equipment. Various voltage outputs (e.g., 400 volts (V), 800V, 1000V, 1200V, etc.) are provided from the hybrid architecture, as well as connections for other accessory or system power (e.g., 28V). The flexible architecture described herein can be quieter than other systems (e.g., quieter than turbine engine systems). For example, noise levels can be less than 70 decibels (dB) at distances of 100 feet or less from current systems.

[0104] The flexible architecture described herein may also be scalable. For example, a larger aircraft may use two or more of the flexible architectures described herein. The flexible architecture may also be used for a variety of aircraft designed for different functions and purposes. For example, the flexible architecture described herein may be useful in Urban Air Mobility (UAM) systems such as electric vertical take-off and landing (eVTOL) aircraft, electric short take-off and landing (eSTOL) aircraft, conventional electric take-off and landing (eCTOL) aircraft, etc. One example of a flexible architecture such as that shown in Figures 17-19 may have the specifications shown in Table 1 below. [Table 1] *Maximum burst shaft power depends on battery configuration. **Dry mass includes engine, generator, inverter, and thermal system.

[0105] As shown above, a 185 kW hybrid unit may be provided, and thus two hybrid units may be provided on a particular aircraft to provide 370 kW of power.

[0106] Figure 20 illustrates a perspective view 1200 of another example of a flexible architecture for an aerospace hybrid system in accordance with an illustrative embodiment. The flexible architecture of Figure 20 includes an engine 1205 and a generator, which are hidden or not visible due to other components of the system, such as cooling ducts. However, similar to the hybrid units of Figures 17-19, a mechanical power output 1210 and an electrical output 1220 (both of which can optionally receive electrical power as well) are provided.

[0107] In this manner, various embodiments herein provide a hybrid power plant that can be integrated into a variety of different types of aircraft in the aerospace market. In doing so, aircraft manufacturers may not need to build their own systems of engines, generators, power electronics, cooling systems, and / or control systems to power the aircraft. This can be advantageous because the development process to create a power generation system and make it aerospace standard can take four or more years and cost more than $10 million.

[0108] In this manner, the hybrid power plant or flexible architecture described herein may be designed, manufactured, etc., separate from the design of the aircraft. Some aspects of the flexible architecture may be customized according to the desires of the aircraft manufacturer, but in a manner that does not result in a redesign or reconfiguration of the entire system. Thus, embodiments herein provide an integrated unit that includes an engine, generator, power electronics, cooling system, and / or control system in one package that is installed on the aircraft. Combining these elements into a single standalone unit may further advantageously allow the unit to go through the Federal Aviation Administration (FAA) certification process as a system. This would allow multiple aircraft manufacturers to use the certification system, reducing the certification burden and development burden on aircraft developers, as well as improving efficiency by eliminating the need for multiple aircraft manufacturers to certify many different power generation systems designed specifically for their aircraft.

[0109] By providing a combined unit with engines, generators, power electronics, cooling systems, and / or control systems, the hybrid flexible architecture described herein can be optimized as an entire system rather than as separate components. Optimize the whole system rather than optimizing the parts. Furthermore, such hybrid units can be used in multiple aircraft designs, but the systems designed as part of the aircraft design process are configured in a way that makes them difficult to reapply elsewhere. Having a hybrid unit that can be applied to multiple market segments and aircraft designs with common power requirements expedites the development of production aircraft, with major components of the aircraft already certified (e.g., hybrid units or flexible architectures, etc.).

[0110] Aviation hybrid electric systems have historically been designed from the ground up for each application / aircraft. Such a process is inefficient and is addressed by embodiments herein. For example, some aircraft are equipped with unique power generation equipment designed specifically for that aircraft. Such solutions may include custom engines, generators, power electronics, control systems, cooling systems, battery packs, propulsion motors, and / or propellers. Embodiments herein provide a compact hybrid system for an aircraft that may constitute two separate halves within an aircraft's power and propulsion system, i.e., one half of the upstream end and one half of the downstream end of a powertrain (such as the hybrid powertrain described herein).

[0111] 21 illustrates examples of downstream components 1305, 1310 and upstream components 1315, 1320 for propelling an aircraft 1300, according to an exemplary embodiment. For example, the downstream components 1305, 1310 of an aircraft system may include motors, rotors / propellers, attitude control components, etc., as may be relevant depending on the particular design of the aircraft. The upstream components 1315, 1320 of an aircraft, which may be repeatable within different aircraft, may include any of engines, generators, batteries, power distribution, fuel, generator noise abatement, etc.

[0112] Specifically, the upstream end of the powertrain may include hybrid powertrain elements responsible for generating electrical power. Such components may include an engine, a generator, power electronics, a control system (for the upstream power generating components), a cooling system (for the upstream components), a battery pack, and / or fuel. The downstream end of the powertrain may include hybrid powertrain elements responsible for converting electrical power into thrust, attitude control, and / or active control of aerodynamics. These downstream components may further include electric motors, propellers, motor controllers, and / or control systems for the propulsion system.

[0113] Thus, there may be a need for a common upstream powertrain across very different electric aircraft designs with similar size and total power requirements. However, because downstream powertrains have little consistency between aircraft, these components may not be standardized to work across many aircraft designs as the upstream components. Furthermore, upstream elements that aid in standardization may include components related to power requirements rather than total energy requirements. In the case of engines, generators, power electronics, cooling systems, and / or control systems, these elements of the upstream powertrain may be sized to fit the specific power requirements (kW or hp) of the aircraft. However, the amount of fuel and the size of the battery pack may be determined by the total energy requirements (kWh or hp hr), which may vary from aircraft to aircraft. In such an embodiment, the volume of fuel may be adjusted by changing the size of the fuel tank to fit the requirements of the aircraft design, and the capacity of the battery pack in kWh may be adjusted by adjusting the number of parallel stacks of cells in the battery pack or by adding additional battery packs.

[0114] Accordingly, embodiments are provided herein for providing a hybrid power plant that tightly integrates an engine, generator, power electronics, control system (for the power plant), and / or cooling system in a weight- and space-efficient manner, such that the hybrid power plant can be certified as a stand-alone unit that is separable from the aircraft and designed to provide propulsion.

[0115] Additionally, as described herein, the rotor in the generator can be optimized to serve multiple purposes in the context of a hybrid power plant. A conventional internal combustion engine may have a flywheel mass attached to the rotating shaft to enhance smoothness of operation. However, in the context of an aerospace system, adding extra mass may be undesirable. As described herein, when an engine is coupled to a generator in a hybrid power plant, the rotor of the generator may be designed to withstand torque impulses from the engine and to be the rotating mass that the engine utilizes to smooth operation.

[0116] Additionally, while auxiliary power units are known in the art, these systems may be designed for purposes other than as the primary propulsion source for an aircraft and therefore may not have a control system that can be certified to standards required for use in propulsion. Additionally, such systems may be designed without a cooling system, which is left to the airframe designer. As such, these systems are not Part 33 certified. Additionally, these auxiliary power unit systems are designed as light-weight auxiliary systems that are used intermittently, rather than as highly efficient propulsion systems used during all phases of flight. Additionally, while auxiliary power units may be designed to generate alternating current (AC) power, the hybrid power plants described herein are capable of generating direct current (DC) power, and therefore the hybrid power plants may be coupled to large propulsion battery packs, as the battery packs are powered and charged using a DC power source.

[0117] Turbogenerators are one type of adaptive auxiliary power unit proposed for hybrid power. Such systems lack cooling system integration that provides airframe developers with a cooling system that is part of the hybrid power plant. As such, airframe developers may need to design their own cooling system associated with the use of a turbogenerator. Using embodiments herein may be advantageous in that a separate cooling system for cooling the hybrid power plant described herein does not need to be designed or developed for a particular airframe, since such a cooling system is already included in the flexible architecture described herein.

[0118] As such, the flexible architectures and hybrid power plants described herein advantageously provide an engine for converting liquid fuel (or gaseous fuel) into rotary mechanical power, a generator coupled to the engine configured to convert the rotary mechanical power into electricity, and / or power electronics coupled to the generator configured to convert a direct AC output of the generator into high voltage DC power. The flexible architectures and hybrid power plants described herein further advantageously provide a control system configured to vary the power output of the engine to match the power demands of the aircraft's main propulsion electric bus to meet the power demands of the aircraft.

[0119] The hybrid powerplant control system, power electronics, generator, and / or engine designs described herein may further comply with regulatory requirements for the reliability of aerospace propulsion systems (e.g., probability of failure is 10 -6or 10 to the power of minus 6). The flexible architectures and hybrid powerplants may further include a control interface that allows the flexible architecture or hybrid powerplant to communicate with a vehicle level flight control system to provide propulsion commands from the vehicle level flight control system to the hybrid powerplant control system, which may also advantageously provide for the hybrid powerplant control system to send status messages back to the vehicle level flight control system (e.g., feedback for use in controlling the flexible architecture or hybrid powerplant). The flexible architectures and hybrid powerplants may further include a cooling system that maintains temperature ranges of the generator, power electronics, and / or engine throughout the full range of operating power outputs of the flexible architectures and hybrid powerplants described herein.

[0120] Various embodiments of the flexible architecture or hybrid power plants described herein may further include a control system that varies the power output by varying the engine torque and / or maintains a substantially constant revolutions per minute (RPM) over a significant range of the power output. Such embodiments may provide faster response of the flexible architecture or hybrid power plant by eliminating throttle lag and longer response times associated with the rotational inertia of the system.

[0121] Various embodiments of flexible architecture or hybrid power plants described herein may further include an option to provide a portion of the engine's power output as mechanical shaft power and a portion as DC power. Various embodiments of flexible architecture or hybrid power plants described herein may further include embodiments in which the engine may be a piston engine, a diesel piston engine, a turbine engine, a rotary engine, or other form of combustion engine. Various embodiments of flexible architecture or hybrid power plants described herein may further include examples in which the rotor of the generator is designed to be the flywheel of the engine. Various embodiments of flexible architecture or hybrid power plants described herein may further include a clutch between the engine and the generator so that the generator can be operated as a motor while the engine is shut down in some types of parallel hybrid installations described herein.

[0122] Air Cooling Element As further described below with respect to Figures 22-34, various embodiments described herein also provide for simultaneous air cooling of multiple elements of a hybrid power plant, such as the hybrid flexible architectures described herein. For example, the engine (e.g., a piston engine, a rotary engine, a turbine engine, etc.), the electric machine (e.g., a generator, a motor, or a generator / motor as described herein), the power electronics, and / or the engine intake of a hybrid system may all be advantageously efficiently and simultaneously cooled by the cooling systems described herein. In this manner, disparate components of a hybrid power plant with separate cooling components may be linked with a combined air cooling system that may reduce aircraft weight and increase aircraft reliability.

[0123] Various embodiments of the cooling system described herein utilize air cooling so that air is supplied to different aspects or components of the hybrid power plant. Air is lighter than other media used for cooling, such as water. Thus, the embodiments described herein may have a weight advantage over other systems, such as systems that use liquids, such as water, as the primary medium for cooling. In addition to being heavier than air-based systems, water-cooled systems may face icing issues, especially in aircraft that may operate at high altitudes and thus be exposed to low temperatures.

[0124] Exemplary embodiments advantageously connect a fan, impeller, and / or blower to a power shaft or crankshaft of a flexible architecture described herein (e.g., power shaft 111 of FIG. 9A , crankshaft 160 of FIG. 9B , power shaft 180 of FIG. 9B ), whereby the fan, impeller, and / or blower are mechanically driven based on power provided to the power shaft or crankshaft by an engine of a flexible architecture described herein (e.g., engine 105 of FIG. 9A , engine 155 of FIG. 9B ) or generator / motor (e.g., generator / motor 121 of FIG. 9A , generator / motor 185 of FIG. 9B ). As such, fans, impellers, and / or blowers are configured to provide air cooling directly from mechanical power received from a rotating power shaft and / or crankshaft, and may provide air to multiple system elements for cooling, including providing air directly to the components for cooling or to one or more heat exchangers or finned heat sinks that are used to cool other components (e.g., components with their own liquid cooling systems).Unless otherwise noted, it should be understood that the terms fan, blower, and / or impeller may be used individually to refer to any of fans, blowers, impellers, or other similar components, and any combination of such elements.

[0125] The embodiments described herein provide a lighter weight system than systems that use separate cooling for individual components of the flexible architecture. Additionally, the embodiments herein can reduce conversion losses that may occur in systems that convert mechanical power to electrical power to drive electric fans, since mechanical power from the power shaft or crankshaft can be provided directly to drive the fan. Thus, mechanical power from the flexible architecture is directly converted to air cooling flow. The embodiments described herein further provide a lighter weight and more efficient system, since the cooling fan and associated ductwork can be tightly coupled or located relative to the rest of the flexible architecture, thereby resulting in an efficient, lightweight and compact system for powering an aircraft. The embodiments also increase efficiency by reducing the distance between the cooling inlet of the air cooling system and the device or component being cooled.

[0126] 22 illustrates a schematic diagram of an example cooling system for a hybrid power plant according to an exemplary embodiment. The hybrid power plant may be any of the flexible architectures described and / or illustrated herein, for example, as shown in and discussed with respect to FIGS. 1-8.

[0127] The cooling system of FIG. 22 includes a blower 902 that is powered directly via mechanical energy from a shaft passing through a generator / motor 914. The shaft may also be connected to an engine 904. In this manner, the shaft may be driven by one or both of the generator / motor 914 and the engine 904. The engine 904 may be a piston engine, a turbine engine, a rotary engine, or other type of combustion engine or other engine. An inverter 912 may further be attached to the generator / motor 914, whereby electrical power may be generated from the rotation of the shaft, or electrical power (e.g., from a battery pack or other power source) may be input to the generator / motor 914 and used to rotate the shaft. The engine 904 may further include a cylinder 906, oil for an oil cooling system 908, and a turbocharger 920. An air charge cooler 918 may further be included in the system to cooperate with the turbocharger 920. The system further includes an oil cooler 916 and various hardware 911 (e.g., control or other electronics).

[0128] The blower 902 is configured to be rotated by an engine 904 and / or a generator / motor 914 which rotates a shaft to which the blower 902 is connected. Cool air from the blower 902 may be channeled through various ducting structures to the motor / generator 914, various hardware 911, the cylinder 906 of the engine 904, an oil cooler 916 (e.g., a heat exchanger), a charge air cooler 918, or other components requiring cooling. In various embodiments, some of the components to which the air is channeled may be or include a heat exchanger (e.g., an air-to-air heat exchanger, an air-to-fluid heat exchanger), whereby the air from the blower 902 may be used to indirectly cool the component via the heat exchanger. In various embodiments, any of the components of FIG. 22 or part of the flexible architecture may include a heat sink element, such as a set of fins, configured to sink heat from the component to the air from the blower 902. As such, the components may be indirectly cooled via a heat sink mechanism, which is in contact with the cooling air from the blower 902. In various embodiments, a combination of a heat exchanger and a heat sink (e.g., fins) may be used to cool the components. For example, the heat sink element may reject heat to the air or fluid on a first side of the heat exchanger, and air from the blower 902 may be directed to a second side of the heat exchanger to remove heat from the air or fluid on the first side of the heat exchanger.

[0129] Thus, the blower 902 can be used to cool various components of the flexible architecture, as described further herein. For example, air from the blower 902 can be directed to an oil cooler 916, which is an air-to-fluid heat exchanger configured to exchange heat between the air from the blower 902 and oil in the oil cooler 916. The cooled oil from the oil cooler 916 can then be circulated into an oil cooling system 908 of the engine 904 to cool the engine 904 (e.g., remove heat from the engine 904 by transferring heat to the oil). The hot oil from the oil cooling system 908 can then be circulated back to the oil cooler 916 to be cooled again via the air from the blower 902.

[0130] Cool air may also be provided to the charge air cooler 918. Ambient air may enter the turbocharger 920, be compressed, and then output to the charge air cooler 918. Compressed air from the compressor inlet side of the turbocharger 920 may then be cooled in the charge air cooler 918 using air channeled from the blower 902 to the charge air cooler 918. In other words, the charge air cooler 918 may function as an air-to-air heat exchanger. The cool air may then be output from the charge air cooler 918 to the intake of the engine 904, for use in the combustion cycle of the engine 904, for example. The exhaust output from the engine 904 may then be channeled to the hot side of the turbine or turbocharger 920, which then discharges the air to the environment as exhaust. In this manner, the air used by the turbocharger and / or engine may ultimately be indirectly cooled using the air-to-air heat exchanger of the charge air cooler as part of the turbocharger cycle.

[0131] In this way, various components of a flexible architecture as described herein can be cooled. The cylinders (or rotors) of a diesel aircraft engine (e.g., a piston combustion engine) can be air-cooled or liquid-cooled. In the example of FIG. 22, the cylinder 906 is air-cooled. However, the cylinders may additionally or alternatively be liquid-cooled by adding a heat exchanger between the liquid of the cylinder liquid cooling system and the cold air provided by the blower 902. When a liquid coolant is used, the liquid may be, for example, a mixture of water and glycol. Similarly, the cylinder heads of an aircraft diesel engine (e.g., a piston combustion engine) can be air-cooled, oil-cooled, or water-glycol-cooled. Thus, the air from the blower 902 can be used directly or indirectly (using a heat exchanger or a finned heat sink) as well as the cylinders described herein. Other engines other than piston engines, such as turbine or rotary engines, may also include liquid or air cooling systems for their components and therefore may similarly benefit from the cooling systems described herein (e.g., through direct cooling via air from the blower 902, or a heat exchanger between the air from the blower 902 and a coolant of the engine or a separate cooling system associated with the engine).

[0132] Engine oil in an engine can also be cooled with the flexible architecture. In the example of Fig. 22, oil in the oil cooling system 908 is circulated through an oil cooler 916, with heat exchange occurring between the oil in the oil cooling system 908 and the cool air provided by the blower 902. Heat absorbed by the oil in the oil cooling system 908 in the engine 904 can come from bearing shear in the engine 904, and the oil can also be used for other cooling such as the cylinder head and / or piston (or rotor).

[0133] The charge (intake air) is typically air-cooled, which is required for turbocharging. Turbocharging is very common in aircraft to increase the range of altitudes that can be used with the power to handle the mission, and it also greatly improves the overall thermal efficiency of the engine. Compressing the intake air increases its temperature, and this temperature must be reduced before it is introduced into the cylinder to avoid problems associated with piston cooling, detonation, etc.

[0134] An electric motor / generator (also referred to herein as an electric machine), such as motor / generator 914 in Figure 22, may be cooled due to the presence of electrical resistance and current flow in the electrical and electronic components within the motor / generator 914. This cooling may be accomplished by air cooling, such as from blower 902, or liquid cooling, such as via a heat exchanger supplied with air cooling from blower 902. Liquid cooling may be accomplished, for example, via a mixture of water and glycol or a dielectric (non-conductive) fluid.

[0135] An inverter (with associated power electronics), such as inverter 912 in Figure 22, may be cooled by heat generated within the electrical circuitry, such as high speed switches and other hardware therein. Such cooling may be accomplished via air cooling, such as from blower 902, or via liquid cooling, such as via a heat exchanger supplied with cold air from blower 902. Liquid cooling may be accomplished, for example, via a mixture of water and glycol or a dielectric (non-conductive) fluid.

[0136] Other elements of the hybrid power plant described herein may achieve passive cooling. In other words, the cooling requirements of system elements, including but not limited to clutches (if present), couplers, supervisory or other controllers, fan bearings / seals, etc., may be met by the normal use environment without active design features (fans, pumps, radiators) to enhance the cooling provided. In various embodiments, active cooling may be provided to any component of the aircraft as described herein, as needed, by air cooling, such as from blower 902, or liquid cooling, such as by a heat exchanger supplied with cool air from blower 902.

[0137] As discussed above, air or fluid systems may be used to cool various aspects of an aircraft. However, embodiments herein provide a reduced number of fluid cooling systems that may be used within an aircraft to cool various aspects of the aircraft. A fluid cooling system may use one or more pumps to circulate the fluid. Such pumps may be mechanical or electrical. In the case of mechanical pumps, there is weight and complexity associated with the pump. The pumps themselves must also be installed on the aircraft, increasing the weight and complexity of the aircraft. Pumps may also have bearings, seals, and / or plumbing fittings that may leak. If the pumps are powered by electricity, they may be rated for heat transfer and therefore require relatively high power (e.g., 5000 watts (W) or more).

[0138] The fluid system may be designed to accommodate expansion and contraction of the fluid during use, loss of air during system filling, drainage of the system during use or for other reasons, and / or loss of fluid during the design and / or operation of the aircraft. All of these factors can represent engineering complexities and certification challenges that may be advantageously avoided by using the air cooling systems described herein.

[0139] Fluid systems may also have issues with ice formation, such as at temperatures below -35° F. (-35° C.), which may cause the system to fail or become less efficient, or additional components may be added to avoid ice, further increasing the weight and complexity of the cooling system.

[0140] The fluid system may also use some type of heat exchanger. This may be fluid-to-fluid, transferring heat from a hotter fluid to a cooler fluid, or fluid-to-air, transferring heat to air exhausted overboard. In either case, each heat exchanger represents weight and volume (contributing to the overall weight of the flexible architecture / powerplant system and / or aircraft), represents several potential failure points where leaks may occur (at least two, plus spills and exhausts), and often includes welds that carry certain metal fatigue risks. While some heat exchangers may still be used in the embodiments described herein (e.g., to cool engine oil), it may be advantageous to reduce the number of heat exchangers and / or fluid cooling systems, as described herein.

[0141] In some example aircraft where the fluid cooling system uses a fluid-to-air cooler, depending on the design of the aircraft and the overall system, such systems may use dedicated fans to move the air to perform the desired heat transfer. Such fans may be electrically driven and therefore may require high powered motors rated for the heat transfer (e.g., 5000 W or more). As discussed above, pumps for the fluid system rated for high power may also be used when considering use in heat transfer applications.

[0142] The use of high power rated pumps and fans can be particularly disadvantageous for aircraft cooling systems. Many pumps, coolers, and fans are heavy, complex, take up a lot of space, and can cause multiple potential points of failure. In order to use electric pumps and / or fans to any degree, adequate power must also be provided to keep the cooling system running. For example, when an aircraft is conducting an extended mission (e.g., several minutes or more), stored energy (e.g., batteries) may not be sufficient to power such pumps and fans, and therefore a generator or other power source is provided. In some cases, such a generator may be an alternator mounted directly on the engine, or may be via a separate generator, or may be via one or more DC-DC converters. In particular, in aircraft featuring distributed electric propulsion with high voltage power for one or more lift or propulsion motors, it may be logical to use DC-DC to convert some of this high voltage power to a lower voltage used by the pumps and fans. However, such components add additional complexity and weight to the cooling system.

[0143] Any additional electrical circuits may have additional connections for power, ground, and control. These connections may be heavy and necessarily require size and rigidity (e.g., minimum bend radius), thus necessitating additional volume around the particular electronic device for safe connection and powering the device. Each device to be powered may also be equipped with short circuit protection components such as fuses or breakers that protect the device but may be resettable for safety reasons. The various electronic devices may also include components that provide safe handling for the service crew and / or in some cases control elements to adjust the function of the device to various parameters of the mission. Such components further increase the weight and complexity of the cooling system.

[0144] When DC-DC conversion is used and the voltage direction is from higher voltage to lower voltage, significant heat can be generated resulting in loss of efficiency and additional system elements requiring active cooling.

[0145] Additionally, copper conductors can be used wherever additional electrically powered equipment is added. Copper is often preferred for carrying electrical current in aircraft. The gauge (diameter of the wire) of copper is determined by a combination of the current being used and the local heat transfer available. Everything associated with the wiring can be heavy: the conductor, insulation, connectors at each end, physical support for the wire to prevent chafing, and / or additional armor applied to the wiring to prevent physical damage. Active heat transfer of the conductors and connectors may not be practical, so the size of the conductors may increase to keep temperatures low, resulting in increased weight. Therefore, it is also desirable to reduce the number of electrically powered equipment to reduce the weight and complexity of the system. Therefore, it is also desirable to reduce the number of components or systems that utilize fluid cooling in the aircraft.

[0146] Figure 23 illustrates an example of a hybrid power plant with a cooling system according to an exemplary embodiment. Figure 24 illustrates a cross-sectional view of an example of a hybrid power plant with the cooling system of Figure 23 according to an exemplary embodiment. Figure 25 illustrates a partial cross-sectional perspective view of an example of a hybrid power plant with the cooling system of Figure 23 according to an exemplary embodiment.

[0147] In particular, Figures 23-25 ​​show together a cooling system that may be used in a hybrid power plant, where the cooling is directly driven by mechanical power, simultaneously providing cool air to various systems of the hybrid power plant, thereby achieving various benefits of reducing fluid cooling systems and electrical systems present in the aircraft. A generator / motor shaft may provide power from an engine 1010 and / or a generator / motor (not shown as it is within a housing / shroud 1014) to fan blades 1020. The shaft 1002 may provide power to mechanical components such as a propulsion mechanism such as a fan or a propeller. The air inlet 1004 receives ambient air, and the fan blades 1020 move the air through a duct structure 1006, 1012, 1013, 1016, 1018 and the shroud 1014 (e.g., annular). In the example of Figures 10-12, the fan blades 1020 are centrifugal blowers that direct air approximately perpendicular or perpendicular to the axis of the fan blades 1020. In various embodiments, axial and / or combination blowers may be used in addition to or in place of centrifugal blowers such as those shown in FIGS.

[0148] Because a blower, such as fan blades 1020, may be mechanically driven from the shaft 1002, there may be little conversion loss and the power consumed may be measurable in terms of the pressure and flow of cooling air provided to cool other components of the system. In contrast, an electric fan requires minimal power consumption, including the conversion of shaft power to electrical power (generation), the conversion of voltage (DC-DC), and the transmission of electrical power (I-DC). 2 R losses), and possibly other losses.

[0149] 10-12 can simultaneously cool multiple devices and systems from a single shaft with airflow directing elements attached thereto, rotating at a set revolutions per minute (RPM). In various embodiments, multiple airflow directing elements (fans or blowers) can be attached to a single shaft, such as shaft 1002. In this manner, different airflow directing elements can direct different amounts of air in different directions, such as at different pressure levels, as described for a particular system or component.

[0150] Various embodiments can provide series or parallel cooling (or both) of various combinations of system components. The system shown in Figures 23-25 ​​cools various components in parallel through different ducts. Air can be introduced from fan blades 1020 into duct structure 1018 (point A in Figure 25) and travel through duct structure 1006 to heat exchanger 1008. Heat exchanger 1008 can be used as charge air cooler 918 or oil cooler 916, for example. In Figures 23-25, both charge air cooler (or other air-induced heat exchanger) and oil cooler can be present, but only one of those components is visible in the views of Figures 23-25, while the other is not visible from the view (although partially visible in Figure 23 as heat exchanger 1030). The heat exchangers 1008 and 1030 may be connected to a ductwork 1018 that includes two separate ducts, only one of which (eg, ductwork 1006) is shown in FIGS.

[0151] Air may be introduced from the fan blades 1020 into the duct structure 1016 (point B in FIG. 25) and pass through the duct structures 1012 and 1013 to cool the cylinders of the engine 1010 (similar to cylinders 906 in FIG. 22). A shroud 1014 may be disposed over or around an electric machine (e.g., a generator motor) and may function as a duct structure to provide cool air for cooling such an electric machine. Air may be introduced from the fan blades 1020 into the shroud 1014 (point C in FIG. 25) and pass through the shroud 1014 (e.g., to cool the motor / generator 914 and / or inverter 912 in FIG. 22). Thus, in FIGS. 23-25, a single centrifugal blower may be shaft driven using power from the engine crankshaft or power shaft to allow cooling air to flow into the blower parallel to the axis of rotation. The air then moves radially outward and is collected by three volutes A, B, and C arranged side-by-side around the blower wheel.

[0152] 25, section A of the volute, furthest from the blower wheel attachment to the hybrid power plant, collects the airflow into an enclosed duct 1018. This duct 1018 is then arranged to feed the high pressure cooling airflow to two aluminum heat exchangers in a V-configuration. One of these coolers may be for the engine oil, while the other may be for the engine intake air.

[0153] Section B of the volute between sections A and C (e.g., the center section) collects the airflow into two ducts diametrically opposed, 180 degrees apart. These two ducts are positioned to provide cooling airflow to the piston engine cylinder. Section C of the volute closest to where the fan blades 1020 are attached to the hybrid power plant is a thin section dedicated to cooling the electric motor and inverter. This airflow is contained within the shroud 1014 and can be forced to flow in parallel through the shroud 1014. The shroud 1014 may include machined aluminum fins therein connected to the electric motor and / or inverter housing to allow for cooling airflow and heat transfer from the electric motor and / or inverter housing to the cooling flow.

[0154] Various embodiments may include multiple centrifugal or radial blower wheels and / or multiple axial fan blades, which may rotate at different RPM if a gearbox is used. These blowers or fans may be connected to one or more ducts that feed air to multiple dedicated radiators (e.g., fluid-to-air heat exchangers or air-to-air heat exchangers), or may be directly connected to components designed to be cooled by airflow (such as our cylinders and our motors / generators).

[0155] In various embodiments, a single rotating shaft may be used with two centrifugal blower elements connected back to back to each other and both attached to the shaft as described herein. In such an embodiment, one side of the hub may drive a larger blower that meets multiple cooling requirements with relatively high pressure rise and high mass flow rates. The other side of the hub drives a smaller blower of the same or different radius, providing different levels of pressure rise and mass flow rates.

[0156] In various embodiments, a device including one or more centrifugal blowers and / or one or more axial fan blade sets may be mechanically driven from the crankshaft and / or power shaft of the hybrid power plant, which can achieve different packaging requirements / footprints of the system and / or can be used to provide different airflows having different pressure rise, mass flow rates, or other engineering parameters as desired for a particular aircraft and its respective propulsion system and cooling needs.

[0157] In various embodiments, the mechanical drive system may include gearing or another type of transmission (e.g., belt, continuously variable transmission (CVT), fluid torque converter) to vary the RPM of the fan system relative to the RPM of the crankshaft or power shaft, rather than rotating at only a single RPM. Such a feature achieves all of the benefits described above by avoiding an electrically driven cooling system, and the gearing adds flexibility to the design of the aerodynamic fan / blower.

[0158] In various embodiments, the ductwork of such a system may be made from various components such as aluminum, composites, three-dimensional (3D) printed materials, etc., or any combination thereof to achieve a lightweight system. The duct material may also be molded into complex curved surfaces, providing aerodynamic efficiency, etc. Composite materials such as carbon fiber and epoxy may be used to achieve weight savings compared to materials such as aluminum or other metals. The duct itself (e.g., between the shaft driver blower or fan and the device requiring cooling) is also carefully designed to help balance the pressure drop and air mass flow rate to the multiple devices on the generator. This may include the shape and size of the duct, constrictions along straight or simple duct sections for engineering purposes.

[0159] In various embodiments, a thermostatic control may also be added to prevent overcooling (e.g., when the ambient air temperature is low). Various embodiments may also include active dampers in the ducts to vary the ratio of air flowing through each duct (e.g., to each component being cooled). The temperature of the component being cooled is also monitored by the controller, and if the component becomes too hot, the controller can adjust the air flow to that component to provide a greater volume and / or pressure of cooling air.

[0160] 26 shows a schematic diagram of a second example cooling system for a hybrid power plant according to an exemplary embodiment. A blower 600 takes in air and supplies it to a charge air cooler (e.g., an engine intercooler) where it is sent through duct 602 to an engine 606 to cool the engine's cylinders. Other air can pass through duct 608 to an engine oil cooler 610.

[0161] 27 shows a schematic diagram of a third example of a cooling system for a hybrid power plant according to an exemplary embodiment. A blower 600 can supply air to a charge air cooler via duct 602 and can separately supply air to duct 604 to cool the cylinders of an engine 606. Other air can pass through duct 608 to an engine oil cooler 610.

[0162] Figure 28 illustrates a schematic diagram of a fourth example cooling system for a hybrid power plant in accordance with an exemplary embodiment. Figure 28 is similar to Figure 26, in that, for example, air is also provided from blower 600 to motor 612 (e.g., an electric machine or motor / generator as described herein).

[0163] 29 illustrates a schematic diagram of a fifth example cooling system for a hybrid power plant in accordance with an exemplary embodiment. FIG 29 is similar to FIG 28, except that additional air is provided to a liquid air cooler 616 via duct 614 to cool the motor / generator and / or other power electronics.

[0164] FIG. 30 illustrates a top view of an example hybrid power plant with a cooling system according to an exemplary embodiment. FIG. 31 illustrates a cross-sectional view of the example hybrid power plant of FIG. 30 along line AA according to an exemplary embodiment. FIG. 32 illustrates a cross-sectional view of the example hybrid power plant of FIG. 30 along line BB according to an exemplary embodiment. FIG. 33 illustrates an alternative view of the example hybrid power plant of FIG. 30 according to an exemplary embodiment, showing details of the engine cooling fins. FIG. 34 illustrates a side view of the example hybrid power plant of FIG. 30 with a cooling system according to an exemplary embodiment.

[0165] In particular, Figures 30-34 show the engine 149 with shaft 113 powering the fan wheel 302, bringing air through the blower inlet 139. The air passes through the upper volute 133, the lower volute 177, the right duct 202, and the left duct 204. The fan wheel 302 may be surrounded by a charge air cooler 159, which receives heated charge air from the turbocharger 153 inlet via duct 119 and outputs cooled charge air via duct 127 to the engine 149, cooled by the air provided by the rotation of the fan wheel 302. Figure 32 further shows the engine intake filter 304 and the engine exhaust 306. Figure 32 also shows the fan wheel fins 308 of the fan wheel 302.

[0166] The motor / generator mount 143 also mounts the motor generator 145 to the engine 149. The right duct 202 and the left duct 204 also feed air to the engine baffling 206 to cool the engine 149. FIG. 33 shows the engine cylinder fins 402 that are used to cool the engine with the cool air from the ducts 202 and 204. FIG. 34 further shows an engine oil cooler 502 that can receive air via the duct 202 to cool the oil for the engine, where the oil is supplied to the engine via a supply line 504 and returned to the cooler via a return line 506. A duct separator 508 can also be used to separate a portion of the duct 202 so that some air is directed to the engine cylinders while other air is directed to the oil cooler 502.

[0167] Direct Current (DC) Bus Elements Various embodiments are described herein for implementing a hybrid electric aircraft. Such aircraft may utilize a high voltage electric bus to distribute power to various components of the aircraft, such as motors for the propulsion of the aircraft. In such hybrid electric aircraft, it may be desirable to stabilize the high voltage electric bus within a certain predetermined voltage range (e.g., near a nominal voltage level) so that the propulsion motors can operate properly. Since various embodiments described herein may use a direct current (DC) bus, in particular, it may be desirable to maintain a desired DC voltage range. Advantageously, various embodiments herein efficiently maintain a desired DC voltage range on the DC bus by directly connecting at least one battery or supercapacitor to the DC bus, and further maintain sufficient charge on the at least one battery or supercapacitor to maintain the desired DC voltage range on the DC bus. Such embodiments may prevent voltage spikes that may damage components of the hybrid electric aircraft or electric aircraft (e.g., electric motors and inverters for propulsion) and avoid voltage spikes or voltage drops that may adversely affect the reliability and / or performance and safety of the aircraft or systems of the aircraft.

[0168] In electric aviation, various embodiments of the overall architecture may include one or more power generating devices (e.g., generators) connected to a high voltage DC bus via a low impedance connection to provide power and energy to the bus. Within the same vehicle, one or more power consuming devices (such as electric motors) connected to the same DC bus may receive power and energy from the DC bus. Various embodiments of the electric aircraft may also include energy storage devices, such as battery packs or capacitors (e.g., supercapacitors), that can receive or transmit power as desired depending on the bus voltage and the battery pack voltage.

[0169] For example, if a high voltage generator is generating DC power directly or is operating through a passive rectifier, the DC voltage generated by the motor may be primarily a function of the rotational speed (RPM) of the shaft that spins the generator. For example, a permanent magnet electric motor may generate a voltage based on the rotational speed (RPM). In many applications, the coupling of voltage and RPM may create motor control issues that limit the value of that electric motor in the system. To further utilize brushless motors that do not use permanent magnets, an external voltage reference may be used to maintain a desired voltage level. A particular problem in the aviation industry is that to ensure flight safety, power consumers (electric motors driving fans, propellers, or other devices) must be precisely controlled over a wide range of flight conditions, conditions that may not match the characteristics of the power consumers (e.g., brushless generators). If the high voltage generator being used rotates slower than expected for any reason, the bus voltage may be lower than desired, and motors on that bus may operate less than expected, which may lead to unsafe or undesirable conditions. If such a high voltage generator is rotating faster than expected, the bus voltage will be high and the motor performance may even exceed expected or desired values. Therefore, in applications where generators and motors share a common bus, it may be desirable to design the generators and motors used accordingly. In electrified aviation, it is desirable to precisely control the motors to provide lift, thrust, aircraft attitude, etc., to the aircraft. Therefore, compared to other non-aviation related implementations, it is desirable to be able to better control the power supplied to the motors (e.g., via a DC bus) by maintaining the power supplied to the motors at a voltage that keeps the motors operating at a desired performance level. Furthermore, the power supplied to the motors may be rapidly adjustable (e.g., providing a pilot or control system that is flexible and can control the motors over a wide range) so that the pilot or control system of the aircraft can control the motors as needed over a wide range of uses. In various embodiments, the inverter may be used to adjust the output voltage of an upstream generator that may be used to supply power to the high voltage bus.The inverter can be used to precisely control the downstream motor under a variety of load conditions.

[0170] Inverters may allow system designers to extend the operating range of any motor and / or generator by controlling the current. For these inverters to function properly, the bus voltage that powers them may be advantageously set and maintained by other means than motor speed (as it is difficult to precisely control the voltage on the bus if only motor speed is used). Maintaining bus voltage is related to capacitance and expected load variations that exist under all system operating conditions. For example, a high voltage bus and power electronics system may become unstable if the load on that bus changes suddenly or the capacitance (which acts like inertia in similar mechanical systems) is too low.

[0171] In various embodiments, the bus voltage may be established and maintained using a battery pack, a capacitor, or any combination thereof. Such devices may add capacitance and / or electrical inertia to the bus and are passive, meaning that the intended function is entirely governed by physics and does not require control and intervention (e.g., by a controller or control system). Supercapacitors (or ultracapacitors) typically do not have a large energy storage capability, but also have the desirable characteristic of high capacitance. Supercapacitors can respond to very rapid fluctuations with enormous power (e.g., energy over time). That is, they can provide stability to the bus against fluctuations that are relatively short in duration, low in amplitude, or the product of these two values ​​is relatively low. Batteries may be desirable because they have a large capacitance for bus stability and can also store high energy. Batteries may not be able to respond to voltage changes as quickly as supercapacitors, as batteries are often more limited in the rate of power usage, especially when charging (where the discharge power capacity is often 10 times or more than the charge capacity). For example, if current needs to be drawn from the bus (e.g., charging a battery) to maintain a desired voltage level, the battery may not be able to absorb that current as quickly as desired in a particular embodiment (depending on the particular characteristics of the selected battery). However, in some embodiments, one or more battery packs may be sufficient to maintain the desired voltage level on the bus.

[0172] Accordingly, various embodiments are described herein that allow independent control of one or more upstream generators and downstream motors by adding a battery pack and / or a supercapacitor bank with appropriate design to maintain a desired voltage on the DC bus. In an architecture where the voltage and capacitance of these storage elements are directly electrically connected to the primary motor control elements on the bus (and not shielded by other switches, chargers, or similar devices), the battery pack and / or supercapacitor bank provide a lightweight and effective anchor or set point for the high voltage DC bus.

[0173] Battery packs on aircraft may be deployed with hybrid power generation systems to support system safety standards applicable to flight articles. Not only are these battery packs and / or supercapacitors selected to provide the required power or energy, but when set to the correct or desired voltage and connected to a high voltage motor controller, the battery packs and / or supercapacitor banks may provide a second valuable benefit of bus stabilization by connecting the battery packs and / or supercapacitor banks directly to the DC bus. Although the battery packs and / or supercapacitor banks may be advantageously selected for a particular aircraft to have a target voltage, the actual voltage on the bus may naturally vary somewhat due to state of charge (SOC) and varying electrical loads. The battery packs and / or supercapacitor banks may also be advantageously selected to ensure that the actual voltage is unlikely to deviate from the desired range. If the actual voltage deviates from the desired range, or is expected to deviate from the desired range, the aircraft controller or the hybrid generator on the aircraft may adjust the power (e.g., torque) provided to the generator, increasing or decreasing the power provided to the DC bus to maintain the voltage in the propeller within the desired range. Additionally, the RPM may be maintained at a constant or relatively constant level or within a predetermined range. Thus, the electrical power provided to the generator or the power output to the power shaft may be adjusted by adjusting the torque output by the engine, rather than by adjusting the RPM of the engine's output. Additionally, it may be desirable to maintain an actual voltage set point, which may vary to remain within a desired tolerance range for operating the electric motor or other components of the aircraft. Additionally, the battery pack may advantageously act as an auxiliary power source to drive the motor or other components of the aircraft in the event of a failure of the generator or other components of the hybrid generator. This may thus provide an increased level of safety and fault tolerance for the system.

[0174] 35 is a diagram of an example system 168 for providing a stable voltage to a direct current (DC) bus, according to an example embodiment. The system 168 includes a hybrid generator 161, which includes a controller 162, an engine 163 connected to a generator 169 by a shaft 164, an inverter 166, and a direct current (DC) bus 167. The engine 163 can provide mechanical (e.g., rotational) power to the generator 169 via the shaft 164, thereby enabling the generator 169 to generate electrical power (e.g., alternating current (AC) power). The AC power from the generator 169 can be converted to DC power by the inverter 166 and provided to the DC bus 167. The inverter 166 can also convert AC power from the DC bus 167 to AC power that can be used by the generator 169 to provide a power output to the shaft (e.g., when the generator 169 functions as a motor to power aircraft components such as propulsion). The controller 162 may control any of the components of the hybrid generator 161 (e.g., controlling the RPM output to the generator 169). The controller 162 may also measure characteristics of the DC bus 167, such as the voltage on the DC bus and / or the current through the DC bus 167.

[0175] The system 168 further includes aircraft components such as inverters 172 and 176 connected to the DC bus 167, electric motors 174 and 178 connected to the inverters 172 and 176, a controller 181, and battery packs 182 and 184. In various embodiments, the aircraft components may have supercapacitors instead of or in addition to the battery packs 182 and 184. In various embodiments, the battery pack or packs and / or supercapacitors may be included as part of the hybrid generator 161 and connected directly to a DC bus in the hybrid generator 161, regardless of whether the aircraft components have separate batteries and / or supercapacitors. While FIG. 35 shows multiple connections extending from the DC bus 167 of the hybrid generator 161 to the aircraft components 171, other configurations are contemplated herein such as a single connection to another bus of the aircraft components 171, or the DC bus 167 itself being part of the aircraft components 171. The controller 181 may be in communication with the control device 162. In this manner, controller 181 can send information to controller 162 about how it is currently controlling / using inverters 172 and 176, electric motors 174 and 178, or how it plans to use these components in the future. Controller 181 can also monitor and measure the status of battery packs 182 and 184 and send information related to their status (e.g., any measurements related to state of charge, voltage, current flowing into or out of the batteries, etc.) to controller 182. In embodiments in which batteries or supercapacitors are included in hybrid generator 161, controller 162 can monitor such components for similar information.

[0176] In various embodiments, fewer, additional, or different elements may be installed on the aircraft than those shown in FIG.

[0177] FIG. 36 is a flow chart illustrating an example method 203 for maintaining a stable DC bus voltage based on a communication from an aircraft-level controller, according to an example embodiment. In operation 209, the controller (e.g., controller 162 of FIG. 35) may receive a communication from an aircraft controller (e.g., controller 181 of FIG. 35) including power consumption or battery status information. The power consumption information may relate to how power is currently being used by, for example, an aircraft inverter or electric motor. The power consumption information (e.g., information about how the controller will increase or decrease the power provided to the motor at a specified time in the future) may also relate to how power is being used by the aircraft inverter or electric motor. The battery status information may include the state of charge, the actual voltage of the system's batteries or supercapacitors, and / or the current flowing into or out of the batteries or supercapacitors.

[0178] Thus, in operation 213, the controller can determine how to adjust the power output of the hybrid generator to maintain the desired voltage range on the DC bus. For example, if the battery charge level is too low and in danger of not being able to maintain the desired voltage, then in operation 217, the controller can send a command to increase the power output of the hybrid generator, which is enough power to charge the battery. In another example, if the aircraft motors are expected to require significantly more power than they are currently using or currently using, then in operation 217, the controller can send a command to increase the power output of the hybrid generator. The power output may also decrease as well. In either case, the controller can adjust this overall power output to the DC bus by modifying the RPM provided by the engine to the generator. Thus, while the battery pack and supercapacitor alleviate the need to adjust the power output of the hybrid generator in real time, some control or adjustment of the RPM, and therefore the output power to the DC bus, may still be desirable in various embodiments, since the battery pack and / or supercapacitor can maintain the DC bus at a desired voltage level.

[0179] 37 is a flow chart illustrating an example method 301 for maintaining a stable DC bus voltage based on measurements by a hybrid generator level controller, according to an example embodiment. Method 301 is similar to method 203, except that method 301 contemplates measurements that may be performed by the hybrid generator controller itself (e.g., controller 162) rather than receiving such measurements or information from another controller (e.g., an overall aircraft system controller, such as controller 181 of FIG. 35).

[0180] In operation 303, the power available at or aspects of the power flowing through the DC bus are measured by the controller. If the DC bus is measurable by the system-wide aircraft controller, operation 303 may be similarly performed by the system-wide aircraft controller. Similarly, in operation 303, the controller may also measure the battery / supercapacitor status (e.g., state of charge, current, voltage, etc.) if the battery and / or supercapacitor is packaged as part of the hybrid generator rather than located as part of the overall aircraft system. In operation 307, the controller determines how the power output of the hybrid generator should be adjusted based on the measurements. For example, if the DC bus voltage is approaching outside of a desired range, in operation 309, it may be desirable to send commands to components of the hybrid generator to adjust the power output of the hybrid generator based on the determination in operation 307 to ensure that the DC bus voltage remains within a desired voltage range.

[0181] FIG. 38 is a diagram of an example of a computing environment including a general-purpose computing system environment 100, such as a desktop computer, laptop, smartphone, tablet, or other similar device capable of executing instructions, such instructions being stored in a non-transitory computer-readable medium. The various computing devices disclosed herein (e.g., controller 162, controller 181, processor / controller 205, aircraft main controller 220, processor / controller 280, or any other computing device in communication with these controllers, which may be part of other components of the aircraft) may be similar to computing system 100 or may include some components of computing system 100. Also, while described and illustrated in the context of a single computing system 100, those skilled in the art will appreciate that the various tasks described below may also be performed in a distributed environment having multiple computing systems 100 linked via a local or wide area network. And, the executable instructions may be associated with and / or executed by one or more of the multiple computing systems 100.

[0182] In its most basic configuration, the computing system environment 100 typically includes at least one processing unit 102 and at least one memory 104, which may be linked via a bus 106. Depending on the exact configuration and type of computing system environment, the memory 104 may be volatile (such as RAM 110), non-volatile (such as ROM 108, flash memory, etc.), or a combination of the two. The computing system environment 100 may have additional features and / or functionality. For example, the computing system environment 100 may also include additional storage devices (removable and / or non-removable), including, but not limited to, magnetic or optical disks, tape drives, and / or flash drives. Such additional memory devices may be made accessible to the computing system environment 100, for example, by a hard disk drive interface 112, a magnetic disk drive interface 114, and / or an optical disk drive interface 116. As will be appreciated, each of these devices are linked to the system bus 306 and enable reading and writing to the hard disk 118, the removable magnetic disk 120, and / or the removable optical disk 122, such as a CD / DVD ROM or other optical media. The drive interfaces and their associated computer readable media enable non-volatile storage of computer readable instructions, data structures, program modules, and other data for the computing system environment 100. Moreover, those skilled in the art will appreciate that other types of computer readable media capable of storing data can be used for this same purpose. Examples of such media devices include, but are not limited to, magnetic cassettes, flash memory cards, digital video disks, Bernoulli cartridges, random access memory, nano drives, memory sticks, other read / write and / or read only memory, and / or other methods or techniques for storing information, such as computer readable instructions, data structures, program modules, or other data. Any such computer storage media may be part of the computing system environment 100.

[0183] A number of program modules may be stored in one or more memory / media devices. For example, a basic input / output system (BIOS) 124, containing the basic routines that help to transfer information between elements within the computing system environment 100, such as during start-up, may be stored in ROM 108. Similarly, the RAM 110, the hard drive 118, and / or the peripheral memory devices may be used to store computer-executable instructions including an operating system 126, one or more application programs 128 (which may include functionality disclosed herein, for example), other program modules 130, and / or program data 122. Additionally, computer-executable instructions may be downloaded into the computing environment 100, as needed, for example, via a network connection.

[0184] An end user may enter commands and information into the computing system environment 100 through input devices such as a keyboard 134 and / or a pointing device 136. Other input devices, not shown, may include a microphone, joystick, game pad, scanner, etc. These and other input devices are typically connected to the processing unit 102 by a peripheral interface 138 coupled to the bus 106. The input devices may be connected directly or indirectly to the processor 102 through an interface, such as, for example, a parallel port, a game port, Firewire, or a universal serial bus (USB). For viewing information from the computing system environment 100, a monitor 140 or other type of display device may be connected to the bus 106 through an interface, such as a video adapter 132. In addition to the monitor 140, the computing system environment 100 may also include other peripheral output devices, not shown, such as speakers and a printer.

[0185] The computing system environment 100 may also utilize logical connections to one or more computing system environments. Communications between the computing system environment 100 and a remote computing system environment may be exchanged through an additional processing device, such as a network router 152 responsible for network routing. Communications with the network router 152 may be performed through a network interface component 154. Thus, it will be understood that within such a network environment, for example the Internet, the World Wide Web, a LAN, or other similar types of wired or wireless networks, program modules depicted relative to the computing system environment 100, or portions thereof, may be stored in the memory storage devices of the computing system environment 100.

[0186] The computing system environment 100 may also include location determination hardware 186 for determining a location of the computing system environment 100. In some cases, the location determination hardware 156 may include, for example, a GPS antenna, an RFID chip or reader, a WiFi antenna, or other computing hardware that may be used to capture or transmit signals that may be used to determine the location of the computing system environment 100.

[0187] Although the present disclosure has described specific embodiments, it will be understood that the claims are not intended to be limited to these embodiments unless expressly set forth in the claims. On the contrary, the present disclosure is intended to cover alternatives, modifications, and equivalents that may be included within the spirit and scope of the present disclosure. Furthermore, in the detailed description of the present disclosure, numerous specific details are set forth in order to provide a thorough understanding of the disclosed embodiments. However, it will be apparent to one skilled in the art that systems and methods consistent with the present disclosure may be practiced without these specific details. In other instances, well-known methods, procedures, components, and circuits have not been described in detail so as not to unnecessarily obscure various aspects of the present disclosure.

[0188] Some portions of the detailed description of this disclosure are presented in terms of procedures, logic blocks, processes, and other symbolic representations of operations on data bits within a computer or digital system memory. These descriptions and representations are the means used by those skilled in the data processing arts to most effectively convey the substance of their work to others skilled in the art. A procedure, logic block, process, or the like, is here, and generally, conceived to be a self-consistent sequence of steps or instructions leading to a desired result. These steps are those requiring physical manipulations of physical quantities. Usually, though not necessarily, these physical manipulations take the form of electrical or magnetic data capable of being stored, transferred, combined, compared, and otherwise manipulated in a computer system or similar electronic computing device. For reasons of convenience and with reference to common usage, such data will be referred to as bits, values, elements, symbols, characters, terms, numbers, or the like in connection with the various presently disclosed embodiments.

[0189] However, it should be noted that these terms should be interpreted as referring to physical operations and quantities, and are merely convenient labels that should be further interpreted in light of the terms commonly used in the art. Unless otherwise indicated, and as will be apparent from the discussion herein, throughout the discussion of the present embodiments, discussions utilizing terms such as "determine", or "output", or "transmit", or "record", or "locate", or "store", or "display", or "receive", or "recognize", or "utilize", or "generate", or "provide", or "access", or "confirm", or "notify", or "deliver" are understood to refer to the operations and processes of a computer system or similar electronic computing device that manipulates and transforms data. Data is represented as physical (electronic) quantities in the registers and memory of the computer system, and is transformed into other data that are also represented as physical quantities in the memory or registers of the computer system, or other such information storage, transmission, or display devices described herein or otherwise understood by those skilled in the art.

[0190] Noise reduction elements Various embodiments are described herein for reducing noise emitted by aircraft parts, such as aircraft power plants or components thereof. Although some embodiments described herein relate to enclosures for aircraft power plants, such as engine cowlings, the various embodiments described herein may be used with aircraft components other than power plants and engines, and may also be used to reduce noise emitted from components other than aircraft (e.g., helicopters, airplanes, vertical take-off and landing (VTOL) aircraft, short take-off and landing (STOL) aircraft, etc.). For example, the embodiments described herein may be implemented on any noise source through or around which air may pass, such as parts of boats, motorcycles, automobiles, other motorized vehicles, or even stationary parts through or around which air passes and generates noise.

[0191] Airflows at the inlets and outlets of noise-generating components, such as aircraft power plants, can be a source of noise because airflows entering or leaving these components can act as a medium through which noise and vibrations can propagate. For example, in an aircraft with a hybrid power plant, the hybrid power plant can include a piston, rotary, or turbine engine that emits noise and has air inlets and outlets through which the noise can propagate. Various embodiments are described herein for designing the geometry of the air inlets and / or outlets to reduce the amount of noise that ultimately radiates from an enclosure having noise-emitting components therein, such as an engine cowling. For example, various embodiments having different air inlets and / or exhausts described herein can be configured to have a desired aspect ratio (e.g., length / width aspect ratio), eliminate line of sight in all directions from the engine (e.g., a noisy combustion engine) on the exterior of a noise-sensitive aircraft, and / or separate interior noise-reflecting surfaces with noise-attenuating materials. Advantageously, embodiments as described herein provide a weight-efficient and effective means of reducing operational noise from noise-producing components, such as aircraft hybrid power plants.

[0192] Noise reduction embodiments described herein may be particularly advantageous for use in certain implementations. For example, some aircraft may be equipped with a hybrid power plant that includes an internal combustion engine (e.g., turbine, rotary, piston) and an electric machine, such as an electric motor / generator. Noise from such a hybrid generator may be generated in and transmitted through the exhaust stream from the combustion engine, and may also escape through the air inlet of the combustion engine. While exhaust stream noise can be minimized with methods such as silencers, noise may also be generated from other components of the hybrid power plant and / or the internal combustion engine, such as throughout the engine core, cooling fans, pumps, and other accessory equipment. Minimizing noise may be difficult because the noise may be generated in multiple places at once (e.g., from multiple sources / components).

[0193] Thus, the embodiments described herein are configured to reduce noise emitted by multiple sources (e.g., multiple power plant or engine parts emitting noise simultaneously). The noise may travel from the noise source to the human ear via a medium such as air. The embodiments described herein include enclosing the noise source (e.g., the power plant of an aircraft) and managing airflow into and out of such enclosure (and subsequently the power plant of the aircraft). The embodiments described herein further provide further noise reduction by adding noise attenuating materials to various portions of the enclosure and by various configurations for reducing noise that may escape from the enclosure (including noise that may escape through the inlet or exhaust of the enclosure). The noise attenuating materials may be noise attenuating foam or any other type of suitable material.

[0194] Such noise-damping materials may further be positioned within the enclosure (e.g., within the air inlets and / or exhausts) in specific orientations and / or geometries to limit noise without interfering with the overall system performance (e.g., without interfering with the airflow or from the power plant). Accordingly, orientations and geometries advantageously sized to avoid introducing undesirable pressure losses into the inlet or outlet cooling airflows (e.g., back pressure into the outlet cooling airflow) are also described herein. In various embodiments, the noise-damping materials used may be selected based on noise-damping properties, resistance to liquids, heat, and / or fire, resistance to moisture, resistance to mold, resistance to corrosion, etc., such that the noise-damping material has properties desired for a particular application.

[0195] Thus, advantageously, the embodiments described herein enable noise emitting components to operate with lower noise characteristics, which may be desirable in hybrid power for aviation, for example.

[0196] As just one example of a system that may advantageously use the systems and methods described herein, hybrid power plants configured to generate electrical and mechanical power for aircraft may emit or generate noise that is desirably minimized. For example, such hybrid power plants may include prime movers such as engines that utilize combustion to generate shaft work / power. The combustion may generate noise, and noise from combustion engines in other applications is often either radiated directly to the environment or mitigated using mufflers or similar methods.

[0197] However, in addition to the noise generated by the prime mover (e.g., an internal combustion engine), a hybrid power plant may include one or more other noise sources, including, but not limited to, (i) fuel injectors opening and closing, (ii) pistons hitting cylinder walls inside a piston engine, (iii) fans knocking through air and / or slotted gaps and generating noise, (iv) fluid pumps (oil, water, fuel, etc.), and / or (v) mechanical vibrations traveling through various parts and pieces.

[0198] The collective noise of these and other components of the hybrid power plant may be referred to herein as ambient noise during operation of the hybrid power plant. The emission and propagation of such ambient noise into the surrounding environment may be significantly reduced using various systems and methods described herein.

[0199] A hybrid power plant may have other features that allow noise emitted from the power plant to be released to the environment. An air intake or cold air inlet may be used for combustion in the engine and / or other cooling tasks of the power plant. A warm or hot air exhaust or outlet may also be released to the atmosphere. Because the airflow speed at the engine intake and exhaust is typically low compared to the speed of sound (e.g., less than 0.3 Mach (Ma)), noise generated by components of the hybrid power plant may travel through the engine intake and / or exhaust airflows. In other words, air entering and exiting the hybrid power plant may transmit sound waves, and the systems and methods described herein advantageously describe intake and / or exhaust geometries and materials that significantly reduce noise emitted from the enclosure (cowling) of the noise-producing components (e.g., of the hybrid power plant, the internal combustion engine, related components of the internal combustion engine, etc.).

[0200] 39 illustrates a side cross-sectional view of an enclosure 103 with noise reduction components according to an exemplary embodiment. The enclosure 103 includes an air inlet 151 (e.g., intake) and an exhaust 157 (e.g., exhaust) that may be fluidly connected such that air can flow from the air inlet 151 to the exhaust 157. Within the enclosure 103 is a cavity 117 for components that may use the air flow, such as a combustion engine (e.g., piston, turbine, rotary). In various embodiments, components within the cavity 117 may block the flow of air between the air inlet 151 and the exhaust 157 when the components are not operating. However, when the components are operating, they may use air and / or move air from the air inlet 151 to the exhaust 157. As such, even if components fill cavity 117 and completely block and / or seal the interior of the enclosure between inlet 151 and outlet 157, inlet 151 and outlet 157 can still be considered to be fluidly connected while the components in the cavity are operating (e.g., while the engine is running). In various other embodiments, the components in cavity 117 may not completely block the fluid path between inlet 151 and outlet 157. In such embodiments, air can still flow through the enclosure even when the components in cavity 117 are not operating.

[0201] As further shown in FIG. 39, the enclosure 103 may be constructed with a variety of side walls 147. In various embodiments, the side walls 147 may be of various configurations or shapes as desired based on the application, the components to be fitted inside the enclosure, if air inlets and exhausts are desired. The side walls 147 may also be configured to allow a desired flow of air through the air inlets 151 and exhausts 157, and then provide sufficient air to the inlets of the components in the cavity 117 while allowing sufficient exhaust from the components in the cavity 117. Examples of enclosures of different shapes are shown in and further described below with respect to FIGS. 40-51. Any or all of the side walls 147 may be coated, covered, formed from, or incorporate a noise dampening material to reduce noise within the enclosure 103 and therefore reduce noise that may escape the enclosure 103.

[0202] The sidewall 147 in Figure 39 is configured with an opening that forms the air inlet 151 and an opening that forms the air outlet 157. In addition to forming the cavity 117, the sidewall 147 also forms a noise reduction chamber 173. Although the noise reduction chamber 173 is shown between the cavity 117 and the air outlet 157 in Figure 39, various embodiments may additionally or alternatively include a noise reduction chamber between the air inlet 151 and the cavity 117 or anywhere else there is airflow within the enclosure 103.

[0203] The noise reduction chamber 173 may include noise attenuation elements such as channels formed by vertically oriented walls within the noise reduction chamber. Examples of such channels are shown in and further described with respect to Figures 41, 42, 47-51. In various embodiments, channels oriented in different directions may be used. For example, in addition to vertically oriented walls, the walls may be oriented horizontally, at any angle, etc. In various embodiments, an individual wall or multiple walls may be shaped to be vertically oriented, horizontally oriented, angled, or any combination thereof at different points on the wall. In such embodiments, the walls may be of any shape so long as the channels between the walls allow airflow to pass through.

[0204] By way of example only, various components 123, 129, 135, and 179, such as aircraft hybrid power plant components, may be mounted or otherwise positioned at different locations within cavity 117. Typically, more or fewer components may be included within cavity 117, and the various components may be in different locations within cavity 117 than shown in FIG.

[0205] Because different components 123, 129, 135, and 179 may be at different locations within cavity 117, they may generate or emit noises that emanate from different locations within the cavity. Thus, it may be difficult to specifically tailor the noise reduction elements to every potential noise source within enclosure 103, as described herein. Thus, noise reduction chamber 173 may attenuate noise or vibrations that propagate within the exhaust air as it passes from noise reduction chamber 173 to exhaust port 157 (or, in present embodiments where a noise attenuation element (e.g., noise reduction chamber) is positioned along the path of the air inlet, may attenuate noise or vibrations that propagate within the incoming air as it travels from the air inlet to components within cavity 117). For example, multiple channels within the noise reduction chamber may be formed of a noise attenuating material such that the noise or vibrations are absorbed by the walls of those channels, thereby reducing the amount of noise or vibration present in the air output of exhaust port 157.

[0206] As shown in FIG. 39, different components 123, 129, 135, and 179 may be located further away or closer to the exhaust port 157 relative to one another based on their placement within the cavity 117. Thus, as shown in FIG. 39, the noise reduction chamber may have an overall length A, and some of the components 123, 129, 135, and 179 may have noise and / or exhaust traveling through at least the length B of the noise reduction chamber 173. Other components of the components 123, 129, 135, and 179 may have noise and / or exhaust traveling further through some or all of the length C of the noise reduction chamber 173 in addition to the length B of the noise reduction chamber 173. Thus, in various embodiments, it may be desirable to extend the noise reduction chamber beyond the cavity 117 (e.g., length B) to ensure that noise emitted from any source within the cavity travels at least a minimum distance B within the noise reduction chamber 173. As such, the noise reduction chamber 173 includes a first section associated with a length C that is immediately adjacent the cavity 117 and a second section associated with a length B that is not immediately adjacent the cavity 117. That is, air from the cavity 117 passes through the first section (e.g., length C) of the noise reduction chamber 173 before passing through the second section (e.g., length B) and exiting through the exhaust port 157.

[0207] The noise reduction chamber 173 may have a height D. The walls within the noise reduction chamber 173 may be configured to have a height approximately equal to D and a length approximately equal to A such that the walls substantially fill the volume of the noise reduction chamber 173 (e.g., as shown in FIG. 42 ). There are openings between the cavity 117 and the noise reduction chamber 173, and also openings in the sidewall 147 of the exhaust port 157, so that air can flow from the cavity 117 between the walls within the noise reduction chamber 173 and out through the exhaust port 157. The walls within the noise reduction chamber 173 may be formed from a noise attenuating material, so that noise in the air moving through the noise reduction chamber 173 may be eliminated or reduced before the air is exhausted through the exhaust port 157. Although a noise reduction chamber associated with the exhaust port has been discussed and illustrated with respect to FIG. 39, it should be understood that a similar noise reduction chamber may be implemented in any space formed by the sidewall between the cavity 117 and the air intake port 151, or even in any space within the cavity 117 itself.

[0208] FIG. 40A is a front perspective view of an exemplary enclosure 207 having an air inlet 211 therein, according to an exemplary embodiment. The enclosure 207 is similar to that shown in FIG. 39, particularly showing the air inlet 211 side of the enclosure 207 and the sidewall 219 that partially forms a cavity within the enclosure 207 configured to hold aircraft hybrid power plant components such as an internal combustion engine and associated components. FIG. 40B is a side view of the exemplary enclosure of FIG. 40A according to an exemplary embodiment. FIG. 40C is a rear view of the exemplary enclosure of FIG. 40A according to an exemplary embodiment. FIG. 40D is a top view of the exemplary enclosure of FIG. 40A according to an exemplary embodiment.

[0209] Figure 41 is a rear perspective view of the exhaust vent 215 of the enclosure 207 of Figure 40A according to an exemplary embodiment. The enclosure 207 shown in Figure 41 shows an exhaust vent similar to the exhaust vent 157 of Figure 39. The edge of the air inlet 211 and another sidewall 219 that partially forms a cavity for the aircraft hybrid power plant are also visible in Figure 41.

[0210] 41 also shows a number of walls 223 that may be formed within the noise reduction chamber 208 of the enclosure 207 (similar to the noise reduction chamber 173 of FIG. 39). A muffler 222 is also shown, which may be disposed within the noise reduction chamber 208 between two of the walls 223 to further reduce noise emitted into the atmosphere. The walls 223 may extend into the noise reduction chamber 208, with an opening at the top of a channel between the walls 223 of the noise reduction chamber 208 allowing air flow between the cavity of the enclosure 207 and the noise reduction chamber 208 (as further shown in FIG. 42).

[0211] FIG. 42 is a top perspective view of a noise reduction channel 224 in the noise reduction chamber 208 of the enclosure 207 of FIG. 40A according to an exemplary embodiment. In particular, the side wall 219 further forms the noise reduction chamber 208, and the walls 223 form the channels 224. One wall 221 may be shaped differently than the other walls 223, for example to accommodate the muffler 222 shown in FIG. 41. The walls 223 may be formed from a noise attenuating material, such as foam, or any other suitable material. The channels 224 may have a width E. However, in various embodiments, the channels 224 may not all have the same width and / or may have variable widths (e.g., may be wider closer to the enclosure cavity and narrower near the exhaust port 215, or may be narrower closer to the enclosure cavity and wider near the exhaust port 215). In embodiments in which the noise reduction chamber has an irregular shape, or any shape other than that shown in Figures 39-42, the walls may have various shapes to fit the noise reduction chamber and may be formed with desired ratios to form a desired channel size (e.g., width E, lengths A, B, C, height D, etc.). The walls 223 may also have widths that vary as desired, or may have a desired width optimized for noise reduction based on a particular application, materials selected, etc. Additionally, each of the channels may have a cross-sectional area through which air flows. That cross-sectional area may be constant or variable over the length of an individual channel, or over the length of the plurality (or all) of the channels. As discussed above, the dimensions of the walls and the spacing between them may vary, so that the cross-sectional area of ​​the channel formed by the walls may also vary. For example, the cross-sectional area may be larger closer to the exhaust port 215 than near the cavity of the enclosure 207, or the cross-sectional area may be smaller closer to the exhaust port 215 than near the cavity of the enclosure 207.

[0212] Thus, the walls within the noise reduction chamber, or other portions of the enclosure or cowling, may be arranged in any manner desired to achieve noise attenuation. Various possible sizes of the walls and their associated channels may be referenced based on different aspect ratios applied to the geometry of the walls and channels. For example, the length / width ratio of only the second section (e.g., the portion of the noise reduction chamber 208 that protrudes to the rear of the cavity) may be length B of FIG. 39 across width E of FIG. 42. By way of example only, a desired length-width ratio may be desirably at least 1.3, e.g., a length B of 12 inches and a width E of 9 inches may be used with an aspect ratio of 1.333. Other aspect ratios may be used to configure the walls and channels of the noise reduction chamber, which may include any of the dimensions A, B, C, D, and / or E as shown in FIG. 39 and FIG. 42.

[0213] These aspect ratios may be advantageously configured to form channels having a desired aspect ratio in terms of allowing adequate airflow through the noise reduction chamber. For example, a low pressure drop passage of air may be desired at either the input or output of the enclosure. On the one hand, if the channel formed by the parallel walls is too wide (e.g., if the foam is spaced too widely compared to the length and height of the channel), the quality of the noise reduction may be compromised. On the other hand, if the channel is narrow and very long, it is quite possible that the noise pressure waves will be attenuated by contacting multiple walls. Thus, for a particular application, type of wall material, etc., a balance of channel width (e.g., length E or distance between two parallel faces), height (e.g., length D), and length (e.g., length A or distance along the axis of the primary direction of airflow) is important to advantageously achieve a balance of the desired noise reduction qualities of the channel without significantly affecting the performance of the engine or other components within the enclosure.

[0214] An example of a sound-damping material that may be used in the embodiments described herein includes melamine open-cell foam made from melamine resin. This foam may feature excellent sound absorption with high flame retardancy and resistance to flame and smoke. For example, open-cell or closed-cell foam may be used and may be formed from a variety of materials such as melamine, cellulose, polyethylene, cotton, other suitable materials, or any combination thereof. The walls described herein may have any desired thickness, and a thickness of 1 inch to 2 inches may be used merely as an example. When different materials are used as the sound-damping material, the thickness may vary based on the properties of the material or combination of materials. The sound-damping materials and / or walls described herein may be lined / coated with another material or may not be lined / coated with other materials. The walls configured to attenuate noise described herein may also be patterned in various ways to reduce resistance to airflow (e.g., smoother patterns) and / or to increase noise attenuation. For example, the materials described herein may be formed to have a smooth surface, an egg-crate surface pattern, a pyramidal surface pattern, a wedge-shaped surface pattern, a hemispherical surface pattern, a corrugated surface pattern, any other pattern, or any combination thereof.

[0215] In the embodiment shown in Figure 42, the foam walls are thick enough so that the walls are self-supporting without having an internal support system within the foam of another type of material. Optionally, other materials can be used within the foam to support it, such as a thin center plate (such as carbon fiber) coated on both sides with a noise-damping material such as foam.

[0216] The walls are further arranged to form channels as described herein to allow air to pass between the parallel or substantially parallel faces of the foam. In this manner, sound pressure waves can be attenuated while minimizing restriction of the central flow of air through the noise reduction chamber into and out of the system. It is therefore desirable to configure the channels between the walls such that aerodynamic resistance (e.g., pressure loss) airflow (e.g., airflow used to cool engine components) is not excessive, which may result in reduced engine performance (including, e.g., cooling system performance, etc.). Using appropriately sized channels can minimize the effects of such reduced performance. In various embodiments, non-parallel walls may additionally or alternatively be used as described herein.

[0217] FIG. 43 is a top perspective view of another exemplary enclosure 500 having noise reduction components therein, according to an exemplary embodiment. FIG. 44 is a side view of the enclosure 500 of FIG. 43, according to an exemplary embodiment. FIG. 45 is a front view of the enclosure 500 of FIG. 43, according to an exemplary embodiment. FIG. 46 is a perspective view of the enclosure 500 of FIG. 43, according to an exemplary embodiment, showing the enclosure being partially transparent. The enclosure 500 of FIGS. 43-46 includes an air inlet 507 and an exhaust (not shown in FIGS. 43-46, but a similar enclosure with an exhaust 1007 is shown in FIGS. 47-51). The enclosure 500 may specifically be a cowling for an engine or hybrid power plant for an aircraft. The enclosure 500 may be designed to be an exterior surface of the aircraft, such that the enclosure 500 is aerodynamic and the air inlet 507 is directed toward the front of the aircraft. As shown in Figure 46, engine components 802 may be within a cavity of enclosure 500, although enclosure 500 is partially transparent. As shown in Figures 47-51 and described further below with respect to Figures 47-51, enclosures similar to enclosure 500 may have multiple noise attenuating walls therein to achieve the noise reduction benefits described herein.

[0218] FIG. 47 is a perspective view of another exemplary enclosure 900 according to an exemplary embodiment, partially transparent, showing the enclosure having noise reducing components therein. FIG. 48 is a top view of the enclosure 900 of FIG. 47 according to an exemplary embodiment. FIG. 49 is a side view of the enclosure 900 of FIG. 47 according to an exemplary embodiment. FIG. 50 is a rear view of the enclosure 900 of FIG. 47 according to an exemplary embodiment. FIG. 51 is a rear view of the enclosure 900 of FIG. 47 according to an exemplary embodiment, similar to FIG. 47, except showing the enclosure as opaque.

[0219] In particular, Figures 47-51 show walls 903 in the enclosure 900 forming channels 1003 for reducing noise emitted and / or generated by the engine parts 802. The enclosure 900 further includes an air inlet 1004 and an exhaust 1007, where air can be used by the engine parts 802 and exhausted through the exhaust 1007 after use. The enclosure 900 may be mounted on an aircraft to power such aircraft. Meanwhile, Figures 47-50 show the enclosure as partially transparent so that the components within the enclosure 900 are apparent, while Figure 51 shows the enclosure as opaque to better show the exhaust 1007. As shown in Figures 47-51, the walls may change shape as a result of the shape of the enclosure itself. Thus, as shown in this example, the walls can be customized to fit any space within the cowling or enclosure to reduce noise emitted by the components therein.

[0220] In an exemplary embodiment, any of the operations described herein may be implemented at least in part as computer readable instructions stored on a computer readable medium or in a memory that, when executed by a processor, can cause a computing device to perform an operation.

[0221] The foregoing description of the exemplary embodiments has been presented for purposes of illustration and description. It is not intended to be exhaustive or to be limited to the precise form disclosed, and modifications and variations may be possible in light of the above teachings or from practice of the disclosed embodiments. It is intended that the scope of the invention be defined by the appended claims and their equivalents.

Claims

1. 1. An energy source for an aircraft, the energy source comprising: An enclosure; The engine, A generator, at least one fuel tank configured to supply fuel to the engine; an electrical connector for outputting electrical power to a propulsion system of the aircraft; the electrical power is generated by the generator and output to at least one electrical component or electrical bus of the aircraft; the engine, the generator, and the at least one fuel tank are each housed within the enclosure; the propulsion system of the aircraft is not within the enclosure of the energy source; Energy source.

2. the energy source is a first energy source; the propulsion system for the aircraft includes a propulsion motor; the aircraft includes a second energy source separate from the first energy source; 2. The energy source of claim 1, wherein the aircraft power plant is flight capable and configured to power the propulsion motors using the second energy source with or without the use of the first energy source.

3. 2. The energy source of claim 1, wherein the electrical power generated by the generator propulsion system passes through wiring or the electrical bus of the aircraft before being supplied to the propulsion system of the aircraft, and the wiring or the electrical bus of the aircraft is not within an enclosure of the energy source.

4. The energy source of claim 1 , wherein the at least one electrical component or the electric bus of the aircraft is connected to at least one of an electric propulsion motor or a propulsion battery of the aircraft.

5. 2. The energy source of claim 1, wherein the energy source is configured to power the same propulsion motors that are powered by one or more batteries of the aircraft, the one or more batteries not contained within the enclosure of the energy source.

6. The energy source of claim 1 , wherein the electrical connector is removably connectable to a corresponding electrical connector of the aircraft.

7. the generator is configured to generate electricity when the engine rotates a shaft connected to a shaft of the generator; The energy source of claim 1 , wherein the engine is a rotary engine, a turbine engine, or a piston combustion engine.

8. a noise reduction chamber contained within the enclosure, the noise reduction chamber including a plurality of channels configured to allow air to pass through the noise reduction chamber; walls of the plurality of channels are configured to absorb noise or vibrations present in the exhaust, and further wherein the walls are formed from a noise attenuating material; or The energy source of claim 1 , wherein the walls within the noise reduction chamber are substantially parallel to one another.

9. The energy source of claim 8 , wherein the noise reduction chamber is disposed in a fluid path between the engine and an exhaust of the enclosure.

10. an air-cooling system contained within the enclosure, the air-cooling system including a fan or blower mechanically driven by power output from the engine; The fan or blower is At least one component of the engine or the generator; or a cooling element including at least one of a heat exchanger or a finned heat sink configured to cool at least one component of the engine or the generator; The energy source of claim 1 , configured to direct air towards

11. the fan or blower is configured to direct air through at least two different air ducts contained within the enclosure; The at least two different air ducts include: A cylinder cooling component of the engine; Engine oil heat exchanger, The generator, or Turbocharger Charge Air Cooler configured to direct air to at least two of the The energy source of claim 10 , wherein each of the cylinder cooling components, the heat exchanger, and / or the charge air cooler are housed within the enclosure.

12. the energy source is configured to be mounted on a wing of the aircraft; or The energy source of claim 1 , wherein the energy source is configured to be mounted to an underside of a wing of the aircraft.

13. The energy source of claim 1 , wherein wiring of the electrical connector passes outside the enclosure of the energy source.

14. further comprising mounting hardware configured to enable non-destructive removal of the energy source from the aircraft; or The energy source of claim 1 , further comprising mounting hardware configured to allow only destructive removal of the energy source from the aircraft.

15. 1. A method for using a first energy source in an aircraft, the method comprising: Mounting the first energy source on the aircraft, the first energy source includes an engine and a generator; the engine and the generator are each housed in an enclosure; the aircraft includes a second energy source configured to power a propulsion motor of the aircraft; the second energy source is not disposed within the enclosure of the first energy source; connecting a first electrical connector of the first energy source to a second electrical connector of the aircraft; outputting electrical power from the generator of the first energy source via the first and second electrical connectors to at least one electrical component or electrical bus of the aircraft; the at least one electrical component or the electrical bus is connected to a propulsion motor of the aircraft, and the propulsion motor of the aircraft is configured to be powered by the second energy source without using the first energy source. method.

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