Patent Application
20230399115
Various types of aircraft may be used to transport goods or people, used as a hobby, etc. Various types of aircraft may identify a range based on their fuel capacity, aircraft type, engine or other propulsion system type, weather conditions, etc. That range may be indicative of how far that aircraft can safely fly in given conditions. Aircraft may not operate safely if it is attempted to fly the aircraft beyond its particular range.
In an embodiment, an energy source for an aircraft includes an enclosure, an engine, an electric generator, and at least one fuel tank configured to provide fuel to the engine; and electrical connectors for outputting power generated by the electric generator to at least one electrical component or electrical bus of the aircraft. The engine, the electric generator, and the at least one fuel tank are each housed within the enclosure.
In an embodiment, a method for using a removable energy source for an aircraft includes mounting the removable energy source to the aircraft. The removable energy source comprises an engine and an electric generator, and the engine and the electric generator are each housed within the enclosure. The method further includes connecting first electrical connectors of the removable energy source to second electrical connectors of the aircraft. The method further includes outputting power from the electric generator of the removable energy source to at least one electrical component or electrical bus of the aircraft.
In an embodiment, a energy source for an aircraft includes an enclosure, an engine, an electric generator, and mounting hardware for attaching the energy source to the aircraft. The energy source for the aircraft further includes electrical connectors for outputting power generated by the electric generator to at least one electrical component or electrical bus of the aircraft. The engine and the electric generator are housed within the enclosure.
Aviation today is undergoing a revolution with the widespread adoption of electrified propulsion. Many vehicles are under development across the globe where the delivery of power to fans/props/rotors used for propulsion, lift, and/or control is via electrical wire rather than mechanical shaft. The use of electricity to transfer power to locations remote from its creation and storage is a design factor for many new designs.
Electrified propulsion relies on three key factors: system voltage, energy, and power. Energy is the total storage capacity (measured in kilowatt-hours or kWh) while power is a measure of the flow of energy (measured in kW). Furthermore, many new-generation aircraft using distributed electric propulsion may rely on direct current (DC) for the storage and transmission of the power and energy.
A common device for storing electrical energy and delivery power is a battery pack. With batteries, the conversion of energy involves chemistry—energy or power are added to the battery pack for storage (resulting in a chemical transformation) and then later extracted from the battery pack in a reverse reaction for use as needed. Another device used for storage of energy or power is supercapacitors or ultracapacitors. While these devices can deliver exceptionally high power levels, the total energy storage is very low for a given product weight so they are seldom chosen for primary energy storage on aircraft.
Battery packs may be comprised of individual cells arranged into modules and further arranged into a battery pack of multiple battery modules or packs. Safety regulations and performance requirements may require the use of battery management systems (to maintain balanced cell voltages throughout multiple charge/discharge cycles), cooling systems, venting in case of fire or other unwanted release of chemical products/gases, and/or safety circuitry. All these support systems may contribute to the total mass of the battery pack(s). In many cases, the specific energy of complete battery packs suitable for safe operation on manned aircraft (measured in watt hours per kilogram (Wh/kg)) is far below that which is may be required by an aircraft design and given mission for an aircraft. In short, current batteries may be too heavy for many aircraft designs today and/or for many missions/flight plans desired to be implemented by aircraft today.
Described herein are hybrid-electric gensets to help address the problem of battery's low specific energy. In particular, the high energy density of liquid fuels may be used in an engine to convert to shaft power, and a generator may be used to convert that power to electricity. In this way, the specific energy of the total system can be much higher than batteries, depending on the amount of fuel on board the aircraft. As described herein, hybrid-electric gensets of various embodiments of the present application may deliver more than six times (6×) the specific energy (energy per unit of mass) of a battery and at a desired power level.
Those hybrid-electric gensets may include components for interface and interaction of physical mounts, wiring, liquid fuel storage, and airflow between the aircraft and the hybrid-electric genset. In some cases, an aircraft may be complete with no design provision for the addition of an internal genset but only a plan for battery storage of energy. As such, it may be desirable to have supplemental energy or power in certain circumstances without having to redesign major aspects of the aircraft. As such, the availability of the range-extending energy pods (REEPs) would be highly useful. The REEPs described herein may perform the same exact function as an external battery: attachment of a physical item and an electrical connection.
As such, described herein are various embodiments for a range-extending energy pod (REEP) for an aircraft. The REEP may include a hybrid-electric powerplant and/or other various desired elements in a single, compact package or enclosure. Within the enclosure, an engine, electric generator, one or more fuel tank(s), etc., and anything else for supplying electrical power to an electrified aircraft may be included. As such, the REEP could be advantageously removably connected to an aircraft whenever additional power is desired. For example, some aircraft may have a particular, limited range. However, the REEP as described herein may be attached to such an aircraft to supply additional electric power and thereby extend the range of the aircraft. In this way, the REEP may act similar to a battery from the perspective of the aircraft, in that the REEP may just plug into the existing electrical system of the aircraft such as through a high voltage bus and supply power to the aircraft as a battery would.
The REEP may also include mounting hardware so that the REEP may be easily mechanically fastened to the aircraft and removed from the aircraft as desired. For example, the REEP may be bolted to an aircraft, and electrical connectors, such as two high-voltage wires, may be plugged in. Then the aircraft may be supplied with significant power or energy beyond what its permanently affixed on-board systems may be able to provide. The electrical power or energy from the REEP may be used to drive propulsion systems (e.g., electric motors configured to turn rotors, propellers, etc.) and/or may be used for other purposes by the aircraft, such as powering other electronics (e.g., accessories) or charging batteries of the aircraft. The embodiments described herein therefore advantageously provide for additional power and energy to be used by aircraft.
Battery-powered electric aircraft may be able to perform missions or flights that conventional aircraft using conventional powerplant solutions cannot perform or would not be permitted to perform. For example, electric aircraft may be able to takeoff or land in smaller spaces that conventionally powered aircraft may not be able to takeoff or land in. Electric aircraft may also be permitted to operate in areas where conventional aircraft would not be permitted to operate, for example due to the large amount of noise created by some conventional aircraft.
However, batteries may be quite heavy and, in many cases, may supply insufficient energy or power to enable a given mission or flight path (e.g., may not supply enough energy or power for certain power intensive tasks like takeoff or landing, or may not supply enough energy for a long enough flight route as desired). As such, described herein are various embodiments for a hybrid-electric powerplant genset that converts liquid fuel to electrical energy and power that may be utilized by an electric aircraft (or any aircraft with electric components). The range extending energy pods (REEPs) described herein may advantageously include any fuel, power conversion, thermal treatment, and wiring to make the REEPs (which may function as a hybrid battery) completely independent of the aircraft fuselage and simple to install. As just one example, a REEP may be attached to an aircraft using four bolts for support and two wires to transfer the electricity. As such, the REEPs described herein may be attached to aircraft that have their own propulsion mechanisms and powerplants, such that those aircraft may fly with or without the REEP attached. Instead, the REEP may provide additional power beyond what the aircraft has on its own without the REEP, for example to extend the range of the aircraft, the speed at which the aircraft may fly, etc.
As just one example, an embodiment may include storage for a volume of fuel that is suitable for delivery of 185 kilowatts (kW) of power at 800 volts direct current (VDC) for 3 hours nonstop. This is approximately 555 kilowatt hours (kWh) of energy. Such an embodiment may have a weight of approximately 450 kg, giving an energy density of over 1200 watt hours per kilogram (Wh/kg). Current battery packs available today for aviation, when considered at the pack level including cooling components and required battery management hardware to maintain safe operation, may only deliver a maximum energy density of approximately 200 Wh/kg. As such, example embodiments described herein may provide at least a sixfold benefit based on energy per unit weight compared to battery systems. The embodiments herein also provide for significant simplicity of use for that significant gain in energy density, as the REEPs described herein may be removably attached to an aircraft, for example using only 4 bolts and a simple electrical connection.
Example embodiments of a REEP may include an integrated hybrid-electric genset (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) electrical power (e.g., including any of the components described herein related to power components, such as under the Direct Current (DC) Bus Elements heading below). The hybrid-electric genset 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). Example embodiments of a REEP may further include one or more storage tanks for liquid fuel, along with components to safely connect this fuel storage to the hybrid-electric genset. Example embodiments of a 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).
Example embodiments of a REEP may further include various physical structures, such as an enclosure (e.g., cowling) for the elements of the REEP, mounting hardware, such as a physical frame for the elements of the REEP, holes for bolts used to mount the REEP, etc. For example, a structural frame of the REEP may be attachable to an aircraft at just four attachment points. In other embodiments different numbers of attachment points may be used. The attachment points may be located on or extend below a bottom of the REEP (e.g., for mounting on top of existing aircraft surfaces), located on or extending above the system (e.g., for mounting under existing aircraft surfaces), or may be configured in any other way as desired to provide mounting points for the REEP to attach to an aircraft. Mounting hardware may also include any aspect of a REEP that is designed to facilitate connection of the REEP to an aircraft. For example, any sort of mechanical structure configured to be attached to an aircraft may be part of the mounting hardware. For example, if a cowling, housing, or enclosure of the REEP is designed to be flush with and welded or otherwise fastened to a surface or portion of an aircraft, the cowling, housing, or enclosure may also be mounting hardware. If a frame on which components of the REEP inside a cowling, housing, or enclosure are mounted is also used to securely attach the REEP to an aircraft (e.g., bolting a portion of the frame of the REEP to a portion of the aircraft), the frame or structural component of the REEP may also be part of the mounting hardware.
An enclosure of the REEP may be or may include an aerodynamic firewall package to control cooling flows, 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 to make them more aerodynamic, more visually appealing, safer, and less noisy. The entire REEP including the enclosure may be attached to an aircraft, for example, with 4 bolts. The attachment points may be arranged with enough lateral and longitudinal spread spacing to provide suitable support for the hybrid-electric genset in terms of strength and stiffness. The total weight of an example REEP may be about 1000 pounds (lbs) in an embodiment, so the attachment hardware may include 4×AN-4 bolts (¼-28) in such an embodiment.
Electrical connectors may be used for the transfer of high-voltage current (energy and power). Depending on voltage, this may be a single pair of wires (positive and negative) or more pairs of wires. For example, one embodiment of a REEP may provide 800 VDC with up to 185 kW of power. Such an embodiment may therefore have 230 amps (A) of DC current and a single wire pair using a 3/0 or 4/0 wire size may be used. In various embodiments, other sizes of wire and number of wires may be used.
Various embodiments may also have electrical or electronic communication between the REEP and a control system of the aircraft, thereby allowing the aircraft systems to influence and/or control the startup or shutdown 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 aircraft and pilot use. As such, additional wiring may be used to connect the REEP and the aircraft for communication and control.
Various embodiments also provide for a REEP that has good aerodynamic performance. Since the REEP is configured to attach to an aircraft, such as an external surface of an aircraft, the REEP advantageously has an aerodynamic profile so as not to affect the flight of the aircraft negatively. In other embodiments, the REEP may fit in a battery enclosure of an aircraft within a fuselage or other aerodynamically designed portion of an aircraft, such that the enclosure may not be aerodynamically designed. Examples of aerodynamic shapes may include enclosures with a rounded external shell to provide low drag when connected to an aircraft. Such a shape (e.g., as shown in
In various embodiments, the REEPs described herein may be used to retrofit an aircraft that may not have been designed to have internal spaces or other configurations/spaces to hold a hybrid powerplant internally. As such, an aircraft without a hybrid powerplant may have one or more hybrid powerplants added in the form of the REEPs described herein. In other words, the REEPs described herein may be removable (e.g., put on and taken off an aircraft for certain missions/uses) or may be used to retrofit an aircraft (e.g., where it is desired to use a REEP more permanently for an aircraft, such that an aircraft designed without provisions for hybridization may still be converted or retrofit to have a hybrid powerplant).
In various embodiments, a REEP may also be connected to an aircraft in a more permanent or non-removable way. For example, a housing of a REEP may be connected to a portion of an aircraft using a more permanent method of affixation, such as welding, riveting, or bonding the housing to the wing, fuselage, or other component of an aircraft. In this way, the more permanent attachment than bolts, for example, make the REEP a more permanent installation on the aircraft. Accordingly, the REEPs described may be a self-contained engine, generator, fuel tank, firewall, and noise mitigation all contained within a housing or enclosure that may be removable from an aircraft by disconnecting mechanical connectors or may be more permanently attached to an aircraft. In any case, the enclosure/housing may wall off the components inside at a distinct location outside the fuselage, so that it advantageously isolates these components from the fuselage, offering risk management advantages (e.g., less fire risk at or near the fuselage). In various embodiments, the REEP accomplishes these and other goals by having various components within a housing or enclosure, where the only components that may pass through the enclosure or housing are wiring for power output and/or control signals. The enclosure or housing may also be an aerodynamic enclosure as described and shown herein, such that the REEP may not create to much drag being outside the fuselage, and can be removable as described herein, or at least placed on an aircraft in a more permanent way without redesigning an existing aircraft to accommodate the REEP. For example, hardware such as bolts, nuts, etc. that may be used to connect a REEP to an aircraft may be considered removable because they may be unfastened non-destructively (e.g., they may be used to put on the REEP to an aircraft more than once). In various embodiments, if a component of the REEP is welded to, riveted to, bonded to etc. an aircraft, those fastening mechanisms may only be destructively removed, and therefore may be used when a more permanent connection of the REEP to the aircraft is desired. In various embodiments, mechanisms that may be removed non-destructively may be used alone, mechanism that may be removed only destructively may be used alone, or mechanisms that may be both destructively and non-destructively removed may be used to fasten or otherwise secure a REEP to an aircraft. In addition, the enclosure of the REEPs described herein also serves as a mechanism to channel and reduce noise that is output by the REEP (e.g., including the noise reduction components described herein). As such, the REEP advantageously offers a package of components that provides power to an aircraft either removably or permanently without having to redesign an aircraft or otherwise add components like noise reduction components beyond attaching the REEP itself.
Air may enter a cooling system 10618 at inlet 10624, which may be formed in an inner wall 10626 separating a main compartment of the enclosure 10602 from a noise reduction chamber 10604 of the enclosure 10602. Similarly, another noise reduction chamber 10606 may be affixed to the back of the enclosure 10602 and again may have a wall separating the noise reduction chamber 10606 from the main compartment of the enclosure 10602. The noise reduction chambers 10604 and/or 10606 may have noise reduction elements (e.g., channels formed by walls of noise attenuating material as described herein) within. In this way, the REEP 10600 may be designed to minimize noise output by the REEP 10600 during use.
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 supplied to the aircraft may be DC power and energy, such as would be supplied by a battery pack. In this way, aircraft that are designed to work off of battery power may be supplied with power from the REEP 10600 without having to change how the battery powered components of the aircraft operate.
The REEP 10600 further includes an air inlet filter 10616 which admits filtered air into the engine 10608 for combustion and to produce power.
At 10910, the REEP may be controlled based on a control signal from the aircraft transmitted through the first electrical connectors and second electrical connectors, which may include connectors for controls wiring. For example, one or more controls wires/connectors may be used to receive, from a controller of the aircraft at a controller of the removable energy source, a throttle control signal or power request signal via the connectors. In this way, 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 in the REEP (e.g., temperature, fuel level, power currently being output, etc.). Other signals may be transmitted from the aircraft to the REEP to control the REEP, such as an on/off signal to power the REEP up or down, whether to output AC or DC power and how much of each, etc.
At 10912, the REEP is powered down such that power is no longer being output to the aircraft from the REEP. At 10914, the electrical connectors of the REEP are disconnected from the electrical connectors of the aircraft. At 10916, the REEP is removed from the aircraft. In this way, the method 10900 demonstrates how a REEP may be attached to an aircraft to supply power to that aircraft, and how that REEP may also be removable from the aircraft. In this way, the REEP may be a temporary and removable power source for an aircraft.
As such, described herein are various embodiments for a REEP that may be affixed to an aircraft (using either components that are non-destructively removable or destructively removable) 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 housing/enclosure of the REEP). Such an aircraft may therefore function with the REEP removed (e.g., the aircraft may have its own built-in or internal energy source while the REEP provides an external, additional energy source for providing electric power to the aircraft). The REEP may, for example, be connected to an electrical bus and/or other components such as wiring in the aircraft, where at least one electric propulsion motor and/or at least one propulsive battery (e.g., a battery used to power a propulsion motor) attached to the electric bus and/or other components. In this way, the REEP may be configured to augment the electrical energy already contained in other parts of the aircraft. As such, the REEP may provide energy to the same propulsive motors that may also be fed by batteries or other energy sources elsewhere on the aircraft than where the REEP is mounted or affixed. The REEP may therefore serve as an external energy source to supplement or augment the power already available to the aircraft from its own or internal energy source. The REEP or external energy source may therefore allow an aircraft to be flown or be flyable with or without the REEP/external energy source attached to the aircraft via a structural connection.
Aircraft typically have custom designed propulsion mechanisms and methods for powering those propulsion mechanisms. In this way, the propulsion mechanisms and power supplied to those propulsion mechanisms can be optimized to provide the amount of propulsion needed for a particular type and size of aircraft, while minimizing weight of the components in the aircraft. In other words, the propulsion mechanisms and power for those propulsion mechanisms are often optimized for a particular type and size of aircrafts such that components of one aircraft could not be easily used in a different types of aircraft drive architectures, such as direct drive aircraft, parallel drive aircraft, and serial drive aircraft.
Described herein are various embodiments for a flexible architecture for an aerospace hybrid system and optimized components thereof. A hybrid system may be or may include a system where fuel is burned in a piston, rotary, turbine, or other engine, and an output of the piston engine may be operatively connected to an electric generator for outputting electric power. The embodiments described herein may include flexible systems that can provide power for many different types of aircraft and propulsion mechanisms. Such systems may advantageously reduce the complexity of designing different types of aircraft, may reduce the costs of manufacturing such systems as less customization allows for economies of scale in mass producing the systems, and ultimately may reduce the complexity of aircraft that use the systems described herein.
The flexible architectures described herein may further be used to provide power to propulsion mechanisms in different ways, either in a same aircraft or in different aircraft. For example, a flexible architecture for providing power to propulsion mechanisms may be able to operate in multiple different modes to provide power to different types of propulsion mechanisms. A first aircraft may utilize one, some, or all of the multiple different modes in which the flexible architecture may operate. A second aircraft may utilize one, some, or all of the multiple different modes, and the modes utilized by the second aircraft may be different than those utilized by the first aircraft.
Therefore, different aircraft may take advantage of different modes of providing power to propulsion mechanisms provided by the flexible architectures described herein. While use of the flexible architectures may be customized in this way, the physical hardware of the flexible architectures may be adapted to use by different aircraft with minimal or no changes to the physical components of the flexible architectures described herein. Instead, the use of different modes in different aircraft may be accomplished largely based on how the components of the flexible architectures are controlled using a processor or controller. As such, computer readable instructions may therefore also be stored on a memory operably coupled to a processor or controller, such that when the instructions are executed by the processor or controller, a computing device that includes the processor or controller may control the various components of the flexible architectures described herein to utilize any possible mode of use desired for a particular implementation, aircraft, flight phase, etc.
Power generation and propulsion systems for aircraft may also utilize various cooling systems to ensure that the various components of an aircraft remain at safe temperatures for operation, as well as maintaining components within temperature ranges where they may operate more efficiently. Further described herein are advantageous cooling systems that leverage various aspects of the hybrid architecture described herein to efficiently cool components of a flexible architecture for providing power to propulsion mechanisms of an aircraft.
Aircraft that have hardware for providing different modes of power to its propulsion mechanisms, may have a variety of components for which it may be desirable to provide cooling. Thus, a single cooling system that efficiently moves air to the different components that enable different modes of power may cut down on weight of the aircraft, as well as power consumption of the cooling systems.
The flexible architecture 101 of
As described above, the engine 105 may be a piston combustion engine. A piston combustion engine may advantageously spin an output rotor or shaft at rotations per minute (RPMs) that may be more desirable for direct output to power a generator and/or a propulsion mechanisms (e.g., a propeller) than other engines. For example, a piston combustion engine may have an output on the order of thousands of RPMs. For example, a piston combustion engine may have an output anywhere from 2200 to 2500 RPM, which may be a desirable RPM for a propeller. In particular, a propeller may be designed to have a size that yields a desired tip speed of the propeller based on the RPM output of the piston combustion engine (e.g., of 2200 to 2500 RPM). Other types of engines, such as a turbine engine, may output rotational power on the order of tens of thousands of RPMs, much higher than a piston combustion engine. Another embodiment may drive the motor/generator at the higher RPM of a turbine engine to benefit the efficiency, power output, or other important factors. In some embodiments, a gear box could be added between the output of a high RPM engine and the other components of
The engine 105 may output rotational power to the clutch 115, which may be controlled to engage or disengage the power shaft 111. In other words, the power shaft 111 may be engaged with the rotational output of the engine 105 by the clutch 115, so that rotational force may be transferred between the engine 105 output and the power shaft 111. When the clutch 115 disengages the output of the engine 105 and the power shaft 111, the power shaft 111 may rotate independently of the output of the engine 105. The clutch 115 may be physically located between the engine 105 and the generator/motor 121, and may even contact the engine 105 and the generator/motor 121 on opposing sides in order to reduce the overall footprint of the flexible architecture. In
The generator/motor 121 may also be engaged or disengaged with the power shaft 111. In other words, the generator/motor 121 may be controlled to switch off such that rotation of the power shaft 111 does not cause the generator/motor 121 to generate electrical power. Similarly, the generator/motor 121 may also be controlled to switch on such that the rotation of the power shaft causes the generator/motor 121 to generate electrical power. The generator/motor 121 is referred to as a generator/motor because it may function as either a generator or a motor. In various embodiments, the generator/motor 121 may be referred to as an electric machine, where an electric machine may be an electric generator, an electric motor, or both.
The flexible architecture further includes an electrical power input and output (I/O) 125 connected to the generator/motor 121. As described further herein, the generator/motor 121 may generate electrical power based on rotation of the power shaft 111 that is output via the electrical power I/O 125 or may receive electrical power via the electrical power I/O 125 that may be used to drive the power shaft 111. Wiring for the electrical power I/O 125 may include more than one wire. In various embodiments, the wiring for inputting electric power into the generator/motor 121 may be the same wiring that is used for outputting electric power out of the generator/motor 121. In various other embodiments, first wiring may be used for input of electric power and different second wiring may be used for output of electric power (so that different wires are used for input and output). In various embodiments, the generator/motor 121 may also have wiring connected thereto that is used for control of the generator/motor 121, to relay sensor or other data about the operation of the generator/motor 121 to a controller, etc.
The generator/motor 121 may also act as a driver for the power shaft 111. Upon receiving electrical power via the electrical power I/O 125 from batteries or some other form of electrical energy storage elsewhere in the system, the generator/motor 121 may impart a rotational force on the power shaft 111 to drive the power shaft 111. This may occur as long as the generator/motor 121 is controlled to be switched on to engage with the power shaft 111. If the generator/motor 121 is controlled to be switched off such that it does not engage with the power shaft 111, the power shaft 111 may not be rotated by the generator/motor 121.
Electrical power output from the electrical power I/O 125 may be used to drive an electric motor for an electric propulsion mechanism (e.g., a propeller). Electrical power output from the electrical power I/O 125 may also be used to power and/or charge other devices on an aircraft or aerospace vehicle. For example, electrical power output from the electrical power I/O 125 may be used to charge one or more batteries. The electrical power output from the electrical power I/O 125 may also be used to power other devices or accessories on an aircraft or aerospace vehicle. Because the electrical power I/O 125 also has an input, the power shaft 111 may be driven by any electrical power received via the electrical power I/O 125, such as power from one or more batteries. The power generated by the generator/motor 121 may be an alternating current (AC) power. That AC power may be converted by power electronics (e.g., a rectifier or inverter) into direct current (DC) power and output to a DC bus. That DC bus may be connected to batteries and/or an electric propulsion mechanism. In this way, the electric propulsion mechanism may be supplied with power via a DC bus. In various embodiments, a motor of the electric propulsion mechanism may use AC power, and the DC power from the DC bus may therefore be converted from DC power to AC power before it is used by the electric propulsion mechanism (e.g., by an inverter). In various embodiments, the AC power generated by a generator may be fed directly to a motor or other device without being converted to DC power and back again. In such embodiments, such AC power may be transmitted via an AC power bus or similar wiring.
Any rotation of the power shaft 111 itself, whether driven by the engine 105 or the generator/motor 121, may also be used to drive one or more propulsion mechanisms. For example, rotation of the power shaft 111 may be used to direct drive a propeller or may be used to power an electric motor that drives a propulsion mechanism. The rotation of the power shaft 111 may also drive a gearbox operably connected to another component, such as one or more propellers, one or more rotors, or other rotating devices for various uses on an aircraft.
An accessory pad 131 may also be coupled to the engine 105, and may include a lower voltage direct current (DC) generator for electrical power that is separate from the generator/motor 121 and the electrical power I/O 125, which may be configured for high voltage and high power I/O. In some embodiments, the generator/motor 121 may also have two different windings and the electrical power I/O 125 may have two different outputs (e.g., high voltage and low voltage). Accessory power may be associated with one of the electrical power I/O 125 outputs in addition to or instead of the accessory pad 131 output. The accessory pad 131 may be used to provide power to devices or accessories on an aircraft or aerospace vehicle that does not require high voltage or current outputs that may be output by the generator/motor 121 at the electrical power I/O 125. A high voltage (HV) of an aircraft may be, for example, 400 volts (V) or 800 V, but may also be anywhere between 50 V to 1200 V. A low voltage (LV) of an aircraft may be 12 V, 14 V, 28 V, or any other voltage below 50 V.
The clutch 175 may be configured to engage the power shaft 180 to translate rotational motion from the crankshaft 160 and the output flange 165 to the power shaft 180. The clutch 175 may be further configured to disengage the power shaft 180 such that the power shaft 180 may rotate independently with respect the crankshaft 160 and the output flange 165. In addition,
In various embodiments, the clutch 175 may be different types of clutches or other mechanisms capable of decoupling the power shaft 180 from the output of the engine 155. For example, the clutch 175 may be a plate style clutch, and may be a dry or wet clutch. Such a plate style clutch may be engaged/disengaged or otherwise controlled mechanically, hydraulically, and/or electrically (e.g., by controllers 205, 220, and/or 280 of
Advantageously, the generator/motor 121 of
The memory 210 may be a computer readable media configured for instructions to be stored thereon. Such instructions may be computer executable code that is executed by the controller 205 to implement the various methods and systems described herein, including the various modes of using the flexible architectures herein and combinations of those modes. The computer code may be written such that the various methods of implementing different modes of the flexible architectures herein are automatically implemented based on various inputs that indicate, for example, a particular flight phase (e.g., landing, takeoff, cruising, etc.). In various embodiments the computer code may be written to implement the various modes herein based on input from a user or pilot of the aircraft or aerospace vehicle, or may be implemented based on a combination of user input and automatic implementation based on non-human inputs (e.g., from sensors on or off the aircraft, based on planned flight plans, etc.) The controller 205 may be powered by a power source on the aircraft or aerospace vehicle, such as the accessory pad 131, one or more batteries, an output of the electrical power I/O 125, a power bus of the aircraft powered by any power source, and/or any other power source available.
The controller 205 may also be in communication with each of the engine 230, the generator/motor 235, the clutch 240, the electrical power I/O 245, the accessory pad 250, and/or the sensor(s) 260. In this way, the components of flexible architectures may be controlled to implement various modes as described herein. In various embodiments, engine 230, the generator/motor 235, the clutch 240, the electrical power I/O 245, and the accessory pad 250 may be similar to or may be the similarly named components shown in and described above with respect to
The sensor(s) 260 may include various sensors for monitoring the different components of the flexible architecture 201. Such sensors may include temperature sensors, tachometers, fluid pressure sensors, voltage sensors, current sensors, state sensors to determine, for example, a current state of the clutch 240, or any other type of sensor. For example, voltage and/or current sensors may be used to inform function and settings of a motor/generator, a state chosen for the clutch, or for adjusting any other component of a system. A state sensor could also indicate a specific mode the flexible architecture is being used in, and the system may receive inputs (e.g., from a pilot, from an automated flight controller), to change the system to a different state or mode for a certain phase of flight that may be upcoming. Other sensors may include a pitot tube for measuring aircraft airspeed, an altimeter for measuring aircraft altitude, and/or a global positioning system (GPS) or similar geographic location sensor for determining a location relative to the ground and/or known/mapped structures.
The components of
In various embodiments, the main aircraft controller 220 may also transmit control signals related to future desired power outputs, future flight phase or flight plan information, etc. In this way, the controller 205 may receive and use information about the expected power demands of the aircraft to determine how to control the aspects of the flexible architecture 201 at both a present moment and in the future. For example, flight plan information may be used to determine when battery power should be used, when batteries should be charged, etc. In another example, if a big demand for power is expected, the controller 205 may ensure that the engine 230 is running at a desired RPM to begin delivering a desired level of power.
In various embodiments, the controller 205 may also be in communication 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 the generator/motor 235, control when the batteries are used to directly power another aspect of the aircraft. However, in other embodiments, the main aircraft controller 220 may be in communication with batteries of the aircraft, and/or may relay information related to the batteries and control thereof to the controller 205. Similarly, if the batteries of the aircraft are controlled with the main aircraft controller 220 rather than the controller 205, the controller 205 may transmit control signals related to the batteries to the main aircraft controller so that the batteries may be controlled as needed or desired with respect to the functioning of the flexible architecture 201.
In various embodiments, the electrical power I/O 245 may include two different outputs (e.g., a high voltage (HV) output and low voltage (LV) output) that are associated with two different windings of the generator/motor 235. As such, two different voltages (e.g., HV and LV) may be output and controlled by the controller 205 and/or the main aircraft controller 220. The electrical power I/O 245 may additionally or alternatively have voltage conversion components (e.g., a DC to DC converter) such that two or more different voltages may be output. In such an embodiment, two different outputs may be achieved without the use of two separate windings. The two different outputs may, for example, be output to different power busses on the aircraft, such as a HV bus and a LV bus. The two outputs of the electrical power I/O 245 may also be separately controlled by the controller 205. As such, the outputs may be turned off (e.g., by letting the power shaft and rotor of the generator spin or freewheel with respect to the rest of the motor/generator by turning off field current of the motor/generator).
In some embodiments, the accessory pad may not be controlled by the controller 205 and/or the main aircraft controller 220. The accessory pad may simply always be on when the engine 230 is operating, or may be controlled separately (e.g., by a manual switch flipped by a user) to control when and how power is supplied to accessories on the aircraft.
In some embodiments, the controller 205 may be in communication with a wireless transceiver that may be on-board an aircraft or aerospace vehicle, so that the controller 205 may communicate with other computing devices not hard-wire connected to the system 200. In this way, instructions or inputs for implementing the various modes for the flexible architectures described herein may also be received from a remote device computing device wirelessly. In other embodiments, the system 200 may only communicate with components on-board the aircraft.
The controller 285 may be in communication with one or more of the propulsion mechanism(s) 255 on the aircraft to control them. The controller 285 may also be in communication with one or more sensor(s) 270 on an aircraft or aerospace vehicle, which may be sensors of the aircraft and sensors of the flexible architecture. In particular, the sensor(s) 260 may also be embedded in any of the components of
In either of
Described below are five specific modes that may be implemented using various embodiments of the flexible architecture described herein (including, e.g., the flexible architectures shown in and described with respect to
In a first mode, which may be referred to herein as a hybrid generator mode, a clutch (e.g., the clutch 115 of
In a second mode, which may be referred to herein as a direct drive engine mode, a clutch (e.g., the clutch 115 of
In a third mode, which may be referred to herein as an augmented thrust mode, a clutch (e.g., the clutch 115 of
In a fourth mode, which may be referred to herein as a direct drive generator/motor mode, a clutch (e.g., the clutch 115 of
In a fifth mode, which may be referred to herein as a split engine power mode, a clutch (e.g., the clutch 115 of
As described herein, any of these five modes (or variations thereof) may be used with the single flexible architecture described herein. In addition, certain modes and or combinations of modes may be beneficial for certain aircraft or aerospace vehicle types, certain propulsion mechanism types, certain flight phases of an aircraft or aerospace vehicle, etc.
For example, in a hybrid electric vertical takeoff and landing (VTOL) aircraft with electric motor driven propellers, the flexible architecture herein may be used solely as a source of electrical power. As such, the flexible architecture may drive the aircraft in the first mode (e.g., the hybrid generator mode) during any portion of a phase of flight in which power must be provided to a power bus of the aircraft or one or more motors of the aircraft.
In another example, in an aircraft with a single, large main pusher propeller (e.g., at the rear of a fuselage of an aircraft) and array of electric motors/propellers (e.g., on a wing of an aircraft) the flexible architecture may be used in the fifth mode (e.g., split engine power mode) during takeoff to supply power mechanically to the main pusher propeller and electrically to the wing-mounted motors.
Alternatively, the flexible architecture described herein may be used to power configurations like those shown in
In another example, an aircraft may be a VTOL aircraft with a gyrocopter style main rotor that may be operated powered or unpowered, and may have forward propulsion motors and propellers mounted on wings. In an embodiment, the flexible architecture may be used entirely in the first mode (e.g., the hybrid generator mode) with electrical power supplied from the electrical power input/output (and the generator/motor) driving a motor coupled to the gyrocopter style main rotor and driving the wing-mounted motors using electrical power. In an embodiment, the aircraft may also be configured with a clutch between the power shaft and the gyrocopter style main rotor such that the flexible architecture may use the second mode (e.g., the direct drive engine mode) or the third mode (e.g., augmented thrust mode) to spin the gyrocopter style main rotor (e.g., to get the gyrocopter style rotor up to speed for takeoff). In such an example, the controller may then cause the flexible architecture to switch to the first mode (e.g., the hybrid generator mode) after the gyrocopter style rotor is up to speed (e.g., switch to the first mode for cruising flight). The fourth mode (e.g., the direct drive generator/motor mode) may again be used in the event of an engine failure to use electrical power to drive the power shaft (and therefore the gyrocopter style rotor) from a power source such as one or more batteries.
Accordingly, described herein are advantageous flexible architectures for aircraft through which a variety of modes for supplying power to propulsion mechanisms may be achieved. While particular aircraft and propulsion mechanism configurations may not utilize each mode described herein that a flexible architecture is capable of, the flexible architectures may still be implemented in different aircraft to achieve different modes. Similarly, while an example of a flexible architecture with five different modes for powering propulsion mechanisms is described in detail herein, other flexible architectures with fewer, more, or different modes for powering propulsion mechanisms are also contemplated herein.
For example, a flexible architecture may not have a clutch as described herein and may still be able to implement various modes described herein where it is desirably to have the engine output coupled to the motor/generator and/or an output power shaft of the system. For example, in the first mode, the engine may rotate a power shaft to cause the generator to generate electricity. In the second mode, the engine may direct drive a mechanical propulsion component, for example, but the engine need not be disengaged from the motor/generator or power shaft because the motor/generator can be turned off or allow the power shaft and rotor of the motor/generator to freewheel within the motor/generator. In the third mode, the engine and motor/generator are used to drive the power shaft, so it would not be desirable to disengage the engine and the motor/generator using a clutch. In the fifth mode, the engine may rotate a power shaft to cause the generator to generate electricity and to cause the power shaft to mechanically power a propulsion mechanism. As such, the power shaft need not be disengaged from the engine output in an aircraft that utilizes any of the first, second, third and/or fifth modes as described above. As such, for an implementation that uses any combination of the first, second, third, and/or fifth modes (and not the fourth mode), a clutch may not be used as the system may have the output of the engine constantly connected to the power shaft in the motor/generator. Such an embodiment may be valuable because clutches may be heavy and/or unreliable.
Referring back to
In other various embodiments, a flywheel may not be used. For example, further embodiments of damping systems and apparatuses are described herein that can damp torque on a power shaft (e.g., the power shaft 111) but do not include a flywheel. Further, in various embodiments, a flywheel and other damping systems or components may be used in combination to damp or smooth out torque applied to a power shaft.
For example, the power shaft or rotor within the generator/motor itself may be rigidly coupled to a crankshaft of the generator/motor. In this way, the crankshaft and rotor together can damp the torque pulses on the power shaft or rotor, and may reduce tangential acceleration due to the torque pulses from an engine. In such embodiments, a clutch may be omitted. As such, a damping system would be internal to the generator/motor, and the footprint and weight of the damping systems may be less than a flywheel or other damping system that may be external to a generator/motor. In particular, the rigid coupling of the power shaft or rotor with the crankshaft may increase the inertia of the power shaft or rotor, such that the additional inertia helps prevent the power shaft from slowing down or otherwise rotating in a manner that would make it more susceptible to acceleration from torque pulses of an engine. In such embodiments, the power shaft or rotor and the crankshaft may function similarly to a flywheel.
In various embodiments, a generator/motor having a static inner portion and a spinning outer portion may be used. This may increase an inertia of the spinning portion and may allow the magnets in the generator/motor to spin and avoid being dislodged by torque spikes. In other words, the magnets may be already spinning in the outer portion and therefore may have a constant stabilizing radial force applied in addition to any tangential inertial force due to torque spike acceleration.
A torque damping system may also be configured as part of the power shaft or rotor that connects the output of the engine to the generator/motor. For example, a hub between the power shaft or rotor of the generator/motor may include a coupling that has torsional spring and/or damping properties. Torsional damping couplings may include an elastomeric component or spring (e.g., made from steel or another metal) that reduces potentially harmful torque impulses from being passed from an engine output to a power shaft or rotor of a generator. Torsional damping couplings may be similar to or may also be referred to as a resonance damping coupling. For example, such torsional damping couplings may reduce an overall system weight and size as opposed to systems that use a flywheel or other large damping system. One or more torsional damping couplings may be installed at any one or more of, within an engine, between an engine and clutch, in the clutch, between the clutch and the generator, and/or within the generator to achieve damping before the power shaft or rotor damages components of the generator itself.
Other ways of damping torque on a power shaft or rotor of a generator may also be used. For example, a magnetic field on a generator may be controlled to pulse it such that it acts upon the power shaft or rotor of the generator to cancel some or all of the torque pulses imparted on the power shaft or rotor by an engine. Such pulses on the field of the generator may be controlled based on a measurement of the torque pulses applied by the engine, and may result in the generator components not being damaged by the diesel engine. For example, the third mode described above where both an engine and a generator/motor apply power to a power shaft, pulses to the power shaft from the generator may both apply power to the power shaft and protect the components of the generator from being damaged. In the other modes described herein, pulses to the power shaft using the generator may be applied whenever the power shaft is being driven in whole in part by the engine. Thus, in order to properly protect the components of the generator in such a method, the pulses applied by the magnetic field of the generator to the power shaft or rotor may be configured to correlate to the torque pulses of the engine to properly counteract those torque pulses.
Further described below are examples of how the flexible architectures described herein may be packaged and/or used in an actual aircraft. For example, certain aircraft may use electric motors to drive propulsion systems, and therefore must have sufficient on-board electrical energy or ways to generate such on-board electrical energy to drive those propulsion systems. In addition, regulations in a given jurisdiction may also require sufficient reserve energy to comply with operational regulations of an aircraft. The flexible architectures described herein may provide such electrical energy for propulsion systems and/or reserve energy such that they systems described herein may work with a variety of electric aircraft. For example, the embodiments herein provide for efficient conversion of jet fuel (or other liquid or gas fuel) to electricity, such that electric aircraft may be powered using widely available fuel sources.
In addition, the hybrid unit may include an integrated cooling system 905 that cools various aspects of the hybrid unit, heat exchangers related to the hybrid unit, or heat sinks such as finned attachments for any aspects of the hybrid unit. A power output 910 may be a power shaft (e.g., the power shaft 111 of
By virtue of the power output 910 and the electrical connectors 915, the hybrid unit of
The hybrid unit of
Accordingly, the hybrid units described herein may be used to power an electric or hybrid electric aircraft, and may offer better power than a battery pack alone would. For example, a hybrid unit as shown in
The flexible architectures described herein may also be scalable. For example, in a larger aircraft, two or more of the flexible architectures described herein may be used. The flexible architectures may also be used in different aircrafts designed for different functions and purposes. For example, the flexible architectures described herein may be useful in urban air mobility (UAM) systems, such as electric vertical takeoff and landing (eVTOL) aircraft, electric short takeoff and landing (eSTOL) aircraft, electric conventional takeoff and landing (eCTOL) aircraft, etc. One example flexible architecture, such as the one shown in
As shown above, a 185 kW hybrid unit may be provided. Accordingly, two hybrid units may be provided in a given aircraft to provide 370 kW of power.
As such, the various embodiments herein provide for a hybrid electric powerplants that may be incorporated into various different types of aircraft in the aerospace market. In doing so, aircraft manufacturers may not have to build their own systems that are made up of an engine, a generator, power electronics, cooling systems, and/or control systems to provide power to those aircraft. This may be advantageous, as a development process to create a powerplant system and certify it to aerospace standards may last 4+ years and may cost more than $10M.
As such, the hybrid powerplants or flexible architectures described herein may be design, manufactured, etc. separably from the design of the aircraft. A few aspects of the flexible architectures may be customized as desired by an aircraft manufacturer, but in a way that does not cause the total system to be redesigned or reconfigured. The embodiments herein therefore provide for an integrated unit that includes the engine, generator, power electronics, cooling systems, and/or control systems in one package to be installed on an aircraft. Combining these elements into a single standalone unit further advantageously allows for that unit to go through the Federal Aviation Administration (FAA) certification process as a system. Then, multiple aircraft manufacturers may use the certified system, removing that certification burden and development burden from the aircraft developer as well as adding efficiencies where multiple aircraft manufacturers will not have to seek certification of many different powerplant systems specifically designed for their aircraft.
By providing a combined unit having an engine, generator, power electronics, cooling systems, and/or control systems, the hybrid flexible architectures described herein may be optimized as a whole system rather than as individual components. entire system rather than optimization of the pieces. Additionally, such a hybrid unit may be used in multiple aircraft designs, whereas systems designed as part of an aircraft design process are configured such that it is difficult to reapply them elsewhere. Having a hybrid unit that may be applied in multiple market segments and aircraft designs with common power requirements leads to faster development of aircraft where a major component (e.g., the hybrid units or flexible architectures) of an aircraft is already certified and in production.
Hybrid electric systems for aviation have historically been designed from scratch for each application/aircraft. Such a process is inefficient and addressed by the embodiments herein. For example, some aircraft have unique powerplants designed specifically for the aircraft. Such a solution may include custom engine, generator, power electronics, control systems, cooling systems, battery pack, propulsion motors, and/or propellers. The embodiment herein provide for a compact hybrid system for an aircraft that may make up one half of two distinct halves within an aircraft power and propulsion system: upstream and downstream ends of a powertrain (such as a hybrid powertrain as described herein).
Specifically, the upstream end of the powertrain may include hybrid powertrain elements responsible for producing electrical power. Such components may include the engine, generator, power electronics, control systems (for the upstream power generation components), cooling systems (for the upstream components), battery pack, and/or fuel. The downstream end of the powertrain may include hybrid powertrain elements responsible for turning the 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.
As such, there may be common upstream powertrain needs across very different electric aircraft designs that are of similar sizes and total power requirements. However, the downstream powertrains may have little consistency from one aircraft to the next and therefore these components may not be standardized to work on many aircraft designs the way the upstream components can. Furthermore, the upstream elements that lend themselves to standardization may include the components that are linked to the power requirements but not the total energy requirements. In the case of the engine, generator, power electronics, cooling systems, and/or control systems, these elements of the upstream powertrain may be sized to fit a specific power requirement (kW or hp) of an aircraft. However, the quantity of fuel and the size of the battery pack may be driven by total energy requirements (kWh or hp hr) and these may vary from aircraft to aircraft. In such embodiments, the volume of fuel may be scaled by changing the size of the fuel tank to match the requirements of the aircraft design, and the capacity of the battery pack in kWh may be scaled by adjusting the number of parallel stacks of cells within a battery pack or by adding additional battery packs.
Therefore, provided herein are embodiments for supplying a hybrid powerplant that tightly integrates the engine, generator, power electronics, control systems (for the power generation system), and/or cooling systems in a weight-efficient and space efficient manner that can be certified as a standalone unit designed to provide propulsive power that is separable from the aircraft.
In addition, as described herein, a rotor inside the generator may be optimized to serve multiple purposes in the context of a hybrid powerplant. Conventional combustion engines may have a flywheel mass attached to the rotational shaft to enhance smoothness of operation. However, in the context of an aerospace system it may be unattractive to add extra mass. When an engine is coupled to a generator in a hybrid powerplant as described herein, the rotor in the generator may be designed to withstand any torque impulses from the engine and it may be designed to be the rotating mass that the engine utilizes for smoothness of operation.
Further, while auxiliary power units are known in the prior art, these systems may be designed for different purposes than as a primary source of propulsion power for an aircraft, and therefore may not have control systems capable of being certified to the standards required for use in propulsion. Additionally, such systems may be designed without the cooling systems, leaving that aspect to the airframe designer. As such, these systems are not certified to Part 33 (FAA regulations for aircraft powerplants). Also, these auxiliary power unit systems are designed to be lightweight auxiliary systems that are used intermittently rather than for highly efficient propulsion systems that are used in all phases of flight. Additionally, auxiliary power units may be designed to produce alternating current (AC) power, whereas hybrid electric powerplants as described herein may produce direct current (DC) power so that the hybrid electric powerplants may be coupled to a large propulsive battery pack, as battery packs provide and are charged using DC power.
Turbogenerators are a type of adapted auxiliary power units that have been proposed for hybrid power. Such systems lack cooling system integration that provides an airframe developer with a cooling system that is part of the hybrid powerplant. As such, airframe developers may be left to design their own cooling systems to accompany use of a turbogenerator. Using the embodiments herein, separate cooling systems for cooling the hybrid powerplants described herein may advantageously not need to be designed or developed for particular airframes, as such cooling systems are already included in the flexible architectures described herein.
As such, the flexible architectures and hybrid electric powerplants described herein advantageously provide an engine that converts liquid fuel (or gaseous fuel) into rotational mechanical power, a generator coupled to the engine that is configured to convert the rotational mechanical power to electricity, and/or power electronics coupled to the generator that are configured to convert the direct AC output of the generator to high voltage DC power. The flexible architectures and hybrid electric powerplants described herein further advantageously provide control systems that are configured to vary the power output of the engine to match the power demand on a main propulsive electrical bus of an aircraft to meet the demands of an aircraft for electric power.
Hybrid powerplant control systems, power electronics, generator, and/or engine designs described herein may further comply with regulatory requirements for the reliability of propulsive aerospace systems (e.g., failure should have a probability of less than 10−6 or ten to the power of negative six). Flexible architectures and hybrid electric powerplants may further include a control interface that enables the flexible architecture or hybrid powerplant to communicate with a vehicle-level flight control systems to enable propulsive power commands to be provided from the vehicle-level flight control systems to the hybrid-powerplant control systems, and also advantageously provide for the hybrid-powerplant control systems to send status messages back to the vehicle-level flight control systems (e.g., feedback for use in controlling the flexible architecture or hybrid powerplant). Flexible architectures and hybrid electric powerplants may further include cooling systems that maintain the temperature range of the generator, power electronics, and/or engine over a full range of operating power output of the flexible architectures and hybrid electric powerplants described herein.
Various embodiments of flexible architectures or hybrid electric powerplants described herein may further include control systems that vary power output by varying engine torque and/or maintain rotations per minute (RPM) substantially constant over a significant range of power output. Such embodiments may provide for faster response of the flexible architectures or hybrid electric powerplants by eliminating throttle lag and a longer response time relating to system rotational inertia.
Various embodiments of flexible architectures or hybrid electric powerplants described herein may further include the option to provide a portion of the engine's power output as mechanical shaft power and a portion provided as DC electrical power. Various embodiments of flexible architectures or hybrid electric powerplants described herein may further include that the engine may be a piston engine, diesel piston engine, turbine engine, rotary engine, or other forms of combustion engine. Various embodiments of flexible architectures or hybrid electric powerplants described herein may further include examples where the rotor of the generator is designed to be a flywheel for the engine. Various embodiments of flexible architectures or hybrid electric powerplants described herein may further include a clutch between the engine and generator to enable operation of the generator as a motor that can be operated while the engine is shut down in some types of parallel hybrid installations as described herein.
As described further below with respect to
Various embodiments of the cooling systems described herein utilize air cooling, such that air is provided to different aspects or components of a hybrid powerplant. Air is lighter than other mediums that may be used for cooling, such as water. Thus, the embodiments described herein may have a weight advantage over other systems, such as those that use liquids such as water as a primary medium for cooling. Water cooling systems, in addition to their greater weight than air-based systems, may also encounter problems with icing, particularly in aircraft that may be operated at higher altitudes and therefore experience low temperatures.
An example embodiment advantageously connects a fan, impeller, and/or blower to a power shaft or crankshaft of the flexible architectures described herein (e.g., power shaft 111 of
The embodiments described herein provide for lighter weight systems than those that use separate cooling for individual components of a flexible architecture. In addition, since mechanical power from a power shaft or crankshaft may be provided directly to drive a fan, the embodiments herein may reduce conversion losses that may occur in systems where mechanical power is converted to electric power to drive electric fans. As such, mechanical power from the flexible architecture may be converted directly to air cooling flow. The embodiments described herein further provide for lightweight and efficient systems because the cooling fan and associated ductwork may be closely coupled or placed with respect to the rest of the flexible architecture, thereby yielding an efficient, lightweight, and compact system for powering an aircraft. The embodiments also increase efficiency by reducing distance between a cooling inlet for the air cooling system and the devices or components that are being cooled.
The cooling system of
The blower 902 is configured to rotate upon the engine 904 and/or the generator/motor 914 turning the shaft to which the blower 902 is connected. The cool air from the blower 902 may be directed through various ductwork to the motor/generator 914, the miscellaneous hardware 911, the cylinders 906 of the engine 904, the oil cooler 916 (e.g., a heat exchanger), the charge air cooler 918 (e.g., a heat exchanger), or any other components that are desired to be cooled. In various embodiments, some of the components to which air are directed may be or may include a heat exchanger (e.g., an air-air heat exchanger, an aid-fluid heat exchanger), such that air from the blower 902 may be used to indirectly cool a component via a heat exchanger. In various embodiments, any of the components of
Accordingly, the blower 902 may be used to cool various components of a flexible architecture as further described herein. For example, air from the blower 902 may be directed to the oil cooler 916, which is an air-fluid heat exchanger that is configured to exchange heat between air from the blower 902 and oil in the oil cooler 916. Cooled oil from the oil cooler 916 may then be circulated into the oil cooling system 908 of the engine 904 to cool the engine 904 (e.g., remove heat from the engine 904 by transferring it to the oil). Hot oil from the oil cooling system 908 may then be circulated back to the oil cooler 916 to again be cooled via air from the blower 902.
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. The compressed air from the compressor inlet side of the turbocharger 920 may then be cooled at the charge air cooler 918 using air directed to the charge air cooler 918 from the blower 902. In other words, the charge air cooler 918 may act as an air-air heat exchanger. Cool air may then be output from the charge air cooler 918 to an intake of the engine 904 to be used, for example, in a combustion cycle of the engine 904. The exhaust output from the engine 904 may then be directed to a turbine or hot side of the turbocharger 920, which then outputs the air as exhaust into the environment. In this way, air used by a turbocharger and or engine may ultimately be cooled indirectly using an air-air heat exchanger of a charge air cooler as part of a turbocharger cycle.
As such, various components of a flexible architecture as described herein may be cooled. Cylinders (or rotors) of diesel aircraft engines (e.g., a piston combustion engine) may be air cooled or liquid cooled. In the example of
Engine oil of an engine may also be cooled in a flexible architecture. In the example of
Charge air (induction air) is typically air cooled and this is required due to turbocharging. Turbocharging is very common on aircraft to expand the usable range of altitude with power to meet the mission, plus turbocharging significantly improves overall thermal efficiency of the engine. Compressing the intake air raises its temperature, and this temp must be reduced before being introduced to the cylinders to avoid problems related to piston cooling, detonation, and others.
An electric motor/generator (also referred to herein as an electric machine), such as the motor/generator 914 of
An inverter (with associated power electronics), such as the inverter 912 of
Other elements of the hybrid powerplants described herein may achieve passive cooling. In other words, cooling requirements for system elements including but not limited to a clutch (if present), couplers, supervisory or other controller(s), fan bearings/seals, etc. may be satisfied by their normal service environment with no active design feature (fan, pump, radiator) to enhance the cooling provided. In various embodiments, as needed, active cooling via air cooling, such as from the blower 902, or liquid cooling, such as via a heat exchanger provided with cool air from the blower 902 may be provided to any component of an aircraft as described herein.
As discussed above, air or fluid systems may be used to cool various aspects of an aircraft. However, the embodiments herein provide for reducing the number of fluid cooling systems that may be used in an aircraft for cooling various aspects of that aircraft. Fluid cooling systems may use one or more pumps in order to circulate fluid. Such a pump may be mechanical or electrical. If it is a mechanical pump, there is weight and complexity related to the pump. The pump itself must also be located on the aircraft, adding weight and complexity to the aircraft. A pump may also have, bearings, seals, and/or plumbing joints that may leak. If a pump is powered by electricity, such a pump may be rated for heat transfer and therefore require relatively high power (e.g., 5000 Watts (W) or more).
Fluid systems may also be designed to accommodate expansion and contraction of the fluid during service, for bleeding air during system fill, for system draining during service or for other reasons, and/or provision for fluid spills in design and/or operations of the aircraft. All of these factors may represent engineering complexity and certification challenges and there may be advantages in avoiding them and using the air cooling systems as described herein.
Fluid systems may also have issues with ice formation, such as at temperatures below −35F (−35C). Thus, systems can fail or be less efficient when ice forms, or additional components to avoid ice may be added, which further adds weight and complexity to a cooling system.
Fluid systems may also use a heat exchanger of some sort. This may be fluid-fluid to transfer the heat of a hotter fluid into a cooler fluid, or it may be fluid-air to transfer the heat to air which is exiting overboard. In any case, each heat exchanger represents weight and volume (which contributes to weight for the flexible architecture/powerplant system and/or the entire aircraft), several potential failure points where leaks can take place (at least two, plus a bleed and a drain), and often includes welding which has specific metal fatigue risks. While some heat exchangers may still be used in the embodiments described herein (e.g., to cool engine oil), reducing the number of heat exchangers and/or fluid cooling systems may be advantageous as described herein.
In some example aircraft where a fluid cooling system uses a fluid-air cooler, depending on the aircraft and the overall system design, such a system may use a dedicated fan to move air and execute the desired heat transfer. Such fans may be electrically driven, which may therefore require high power motors that are rated for heat transfer (e.g., 5000 W or more). As discussed above, pumps for fluid systems may also be used that are rated for high power given their use in a heat transfer application.
The use of high power rated pumps and fans may be particularly disadvantageous for an aircraft cooling system. Numerous pumps, coolers, and/or fans may be heavy, complex, take up a lot of space, and introduce multiple potential points of failure. To the extent electric pumps and/or fans are used, suitable electricity must also be supplied to keep the cooling systems running. If an aircraft, for example, is on an extended mission (e.g., more than a few minutes), stored energy (e.g., batteries) may not be sufficient to provide power to such pumps and fans, and therefore a generator or other power source would be provided. In some cases, such a generator may be an alternator directly attached to an engine, it may be via a separate generator, or via one or more DC-DC converters. In particular, on an aircraft characterized by distributed electric propulsion with high-voltage electrical power intended for one or more lift or propulsion motors, it may be logical to use DC-DC to convert a portion of this high-voltage power to low-voltage for use by pumps and fans. However, such components again add complexity and weight to a cooling system.
Any additional electrical circuitry may have additional connections for power, ground, and control. These connections may be heavy, and necessarily have size and stiffness (e.g., minimum bend radius), therefore taking up additional volume around a given electronic device for safe connection and provision of power to the device. Each powered device may also have short-circuit protection components such as a fuse or breaker, which protects the device but may also be resettable for safety reasons. Various electronic devices may also include components that provide for safe handling of service crew and/or possibly may include a control element to tailor the function of devices for various parameters of a mission. Any such components again add weight and complexity to a cooling system.
If a DC-DC conversion is used, and the direction of voltage is from higher to lower voltage, then significant heat may be generated resulting in lost efficiency and yet another system element which may require active cooling.
In addition, wherever additional electrically powered devices are added, conductors made of copper may be used. Copper is often preferred for carrying of electrical current in aircraft. The gauge of copper (diameter of the wire) is determined by a combination of the current in service and the local heat transfer available. Everything associated with wiring may be heavy: the conductor, the insulation, connectors at each end, physical support of the wire to prevent chafing, and/or additional armor applied to the wiring to prevent physical damage. Since active heat transfer of conductors and connectors may not be practical, size of conductors may be increased to keep temperatures low resulting in higher weight. As such, it is again desirable to reduce the number of electrically powered devices to reduce weight and complexity of a system. Similarly, it is therefore also desirable to reduce the number of components or systems that utilize fluid cooling in an aircraft.
In particular,
Because such blowers as the fan blade 1020 may be mechanically driven from the shaft 1002, there may be no of little conversion losses and the power consumed is may be measurable in cooling air pressure and flow rate provided to cool other components of the system. In contrast, electrical fans may suffer from losses due to conversion of shaft power to electrical power (generation), conversion of voltage (DC-DC), transmission of power (I2R loss), and possibly other losses.
As such, the cooling system depicted in
Various embodiments may provide series or parallel cooling (or both) of various combinations of system components. The system shown in
Air may be introduced to ductwork 1016 (at a point B in
With further reference to
Section B of the volute that is between Section A and C (e.g., the middle section) collects airflow into two ducts spaced 180 degrees apart from each other, diametrically opposed. These two ducts are arranged to feed cooling airflow to cylinders of a piston engine. Section C of the volute that is most proximal to where the fan blade 1020 attaches to the hybrid powerplant, is a thin section dedicated to cooling of an electric motor and inverter. This airflow may be contained in a shroud 1014 and forced to flow parallel through the shroud 1014. The shroud 1014 may include within it machined aluminum fins connected to the electric motor and/or inverter housing for the purpose of permitting flow of cooling air and transfer of heat from the electric motor and/or inverter housing to the cooling flow.
Various embodiments may also include more than one centrifugal or radial blower wheel and/or more than one axial fan blade, and they may spin at different RPMs where a gearbox is used. These blowers or fans may be connected to one or more ducts that feed air to a number of dedicated radiators (e.g., fluid-air or air-air heat exchangers) or directly to components that designed to be cooled by airflow (like our cylinders and our motor/generator).
In various embodiments, a single spinning shaft may be used as described herein with two centrifugal blower elements connected back-to-back to one another both attached to the shaft. In such an embodiment, one side of the hub may drive a larger blower that satisfies multiple cooling requirements with relatively high pressure rise and high mass flow. The other side of the hub may drive a relatively smaller blower with the same or a different radius, a provide a different level of pressure rise and mass flow.
In various embodiments, devices may be mechanically driven off the crankshaft and/or power shaft of the hybrid powerplant, including one or more centrifugal blowers and/or one or more axial fan blade sets. This may achieve different packaging requirements/footprints for the system, and/or may be used to provide different airflows with different pressure rise, mass flow, or other engineering parameters desired for a given aircraft and its respective propulsion system and cooling needs.
In various embodiments, a mechanical drive system may not rotate at only a single RPM, but may include gearing or another style off transmission (e.g., belts, continuously variable transmission (CVT), fluid torque converter) to change the RPM of the fan system relative to the crankshaft or power shaft RPM. With such a feature, all the benefits described of avoiding electrically driven cooling systems would be achieved, and the gearing would add flexibility in aerodynamic fan/blower design.
In various embodiments, the ductwork of such systems may be made from various components, such as aluminum, composite materials, three-dimensional (3D) printed materials, etc., or any combination thereof, to achieve a light weight system, where the ductwork material is also capable of being shaped into complex curved surfaces, provide aerodynamic efficiency, etc. Composites such as carbon fiber and epoxy may also be used to bring about a weight reduction in comparison with materials such as aluminum or other metals. The ducts themselves (e.g., between the shaft-driver blower or fan and the devices that require cooling) may also be carefully designed to help balance the pressure drop and air mass flow rate to multiple devices on the genset. This may include shape and size of ducts, constriction along an otherwise straight or simple duct section for engineering purposes.
In various embodiments, thermostatic control may also be added to prevent overcooling (e.g., when ambient air temperature is low). Various embodiments may also include active dampers in ducts to change ratio of air flowing in each duct (e.g., to each component to be cooled). Temperature of those components that are cooled may also be monitored by a controller so that if a component gets too hot the controller can adjust air flow to that component to provide a greater volume and/or pressure of cooling air.
In particular,
A motor/generator mount 143 also mounts a motor generator 145 to the engine 149. The right duct 202 and left duct 204 also provide air to engine baffling 206 to cool the engine 149.
Described herein are various embodiments for implementing a hybrid-electric aircraft. Such an aircraft may utilize a high voltage electrical bus to distribute power to various components of the aircraft, such as motors for propulsion mechanisms of the aircraft. In such a hybrid-electric aircraft, it may be desirable to stabilize the high voltage electrical bus within a specific, predetermined voltage range (e.g., around a nominal voltage level) so that the propulsion motors may perform adequately. Various embodiments described herein may specifically use a direct current (DC) bus, so maintaining a desired DC voltage range may be desirable. Advantageously, the various embodiments herein provide for efficiently maintaining a desired DC voltage range on a DC bus by connecting at least one battery or supercapacitor directly to the DC bus, and further maintaining a 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 be damaging to components of a hybrid-electric or electric aircraft (e.g., electric motors and inverters for propulsion) and avoid voltage spikes or sags that may negatively impact the reliability and/or performance and safety of the aircraft or systems of the aircraft.
In electrified aviation, various embodiments of an overall architecture may include one or more electric power creation devices (e.g., an electric generator) connected via a low-impedance connection to a high voltage DC bus and feeding electrical power and energy onto that bus. In the same vehicle and attached to that same DC bus may be one or more power consuming devices (e.g., electric motors) that receive electrical power and energy from that DC bus. Various embodiments of electrified aircraft may also include energy storage devices such as battery packs or capacitors (e.g., supercapacitors), which may receive or deliver power as desired depending on bus voltage and battery pack voltage.
If a high-voltage electric generator is directly generating DC power or is operating through a passive rectifier, for example, the DC voltage created by the motor may be a function primarily of motor rotations per minute (RPM) of the shaft rotating the electric generator. A permanent magnet electric motor, for example, may create a voltage based on rotational speed (RPM). For many uses, the coupling of voltage with RPM may create an issue for motor control that limits the value of that electric motor in a system. To gain additional usefulness from a brushless motor without permanent magnets, an external voltage reference may be used to maintain a desired voltage level. A unique problem in aviation is that flight safety requires precise control of power consumers over a wide range of flight conditions (electric motors driving fans, propellers, or other devices) that may not match the characteristics of contributors (such as an electric brushless generator). If a high-voltage generator used is turning slower than expected for any reason, the bus voltage may be lower than desired and any motors on that bus may perform below expectations, which may lead to an unsafe or undesirable condition. If such a high-voltage generator is turning faster than expected, bus voltage may be high and motor performance may again be outside expected or desired values. As such, it may be desirable for applications of generators and motors sharing a common bus to design the generators and motors used accordingly. For electrified aviation, precise control of any motor(s) is desirable to provide lift, thrust, aircraft attitude, etc. for an aircraft. As such, as compared to other, non-aviation related implementations, it is desirable to have better control over a power supplied to any motor(s) (e.g., over the DC bus) by maintaining power supplied to the motor(s) at a voltage that keeps the motor(s) operating at a desired performance level. In addition, the power supplied to the motor(s) may be quickly adjustable so that a pilot or control system of an aircraft may control the motor(s) over a wide range of use as needed (e.g., provide a pilot or control system with a flexible, wide range over which they may control the motor(s)). In various embodiments, inverters may be used to regulate an output voltage of an upstream electric generator(s), which may be used to feed a high voltage bus. Inverters may also be used to precisely control downstream motors under varying load conditions.
Inverters may allow a system designer to expand an operating envelope of any motors and/or generators by controlling current. In order for these inverters to function properly, a bus voltage feeding power to the inverters may advantageously be set and maintained by other methods besides motor RPM (as voltage on a bus may be difficult to control precisely where only motor RPM is used). The maintenance of the bus voltage relates to capacitance and the expected variations in load present under all system operating conditions. If that bus has loads that are varying too rapidly or capacitance (which acts like inertia in an analogous mechanical system) that is too low, for example, then the high voltage bus and power electronic system may become unstable.
In various embodiments, bus voltage may be established and maintained using battery pack(s), capacitor(s), or any combination thereof. Such devices may add capacitance and/or electrical inertia to the bus and are passive, meaning their intended function is ruled completely by physics and may not require control or intervention (e.g., by a controller or control system). Supercapacitors (or ultracapacitors) additionally have a desirable feature of high capacitance, though they typically lack significant energy storage. Supercapacitors may respond to very rapid fluctuations with enormous power (e.g., energy over time). In short, they may provide stability to a bus for fluctuations that are relatively short in duration, low in amplitude, or where the product of those two values is relatively low. Batteries may also be desirable because they have significant capacitance for bus stability and may also store high energy. Batteries may not be able to respond to a change in voltage as quickly as a supercapacitor, as batteries often have more limited rate of power applications, particularly in charging (where discharging power capacity is often 10X or more higher than charging capacity). For example, if it is necessary to pull current off a bus to maintain a desired voltage level (e.g., charge a battery), a battery may not absorb that current as quickly as would be desired in certain embodiments (depending on the specific characteristics of a selected battery). In some embodiments, however, one or more battery packs alone may be sufficient to maintain a desired voltage level on a bus.
Accordingly, various embodiments are described herein that enable independent control of one or multiple upstream electric generators and downstream motors by adding a battery pack and/or supercapacitor bank with an appropriate design to maintain a desired voltage on a DC bus. With an architecture where the voltage and capacitance of those storage elements are directly electrically connected to the main motor control elements on the bus (and not shielded by other switches, chargers, or like devices), the battery pack and/or supercapacitor bank provide a lightweight and effective anchor or setpoint for a high voltage DC bus.
A battery pack in an aircraft may be deployed along with a hybrid-electric generation system to support system safety standards applied to flight articles. If these battery packs and/or supercapacitors are chosen not only to provide required power or energy but are also set at a correct or desired voltage and are connected to high voltage motor controllers, the battery pack and/or supercapacitor bank may provide a second and valuable benefit of bus stabilization by connecting the battery pack and/or supercapacitor bank directly to a DC bus. The battery pack and/or supercapacitor bank may also be advantageously chosen for a given aircraft such that it has a target voltage, though actual voltage on the bus may naturally fluctuate some with state-of-charge (SOC) and varying electric loads. The battery pack and/or supercapacitor bank may also be advantageously chosen so that the actual voltage is unlikely to go outside of a desired range. In instances where the actual voltage does go out of the desired range or is expected to go out of the desired range, a controller of the aircraft or a hybrid-electric genset in the aircraft may adjust the power (e.g., torque) supplied to the generator to add or reduce electric power supplied to the DC bus to maintain the voltage within a proper, desired range. RPM may further be maintained at a constant or relatively constant level or within a predetermined range. Therefore, power supplied to the generator or otherwise output to a power shaft may be adjusted by adjusting the torque output by the engine rather than through adjustment of the RPM of the output of the engine. It may further be desirable to maintain an actual voltage set point that may fluctuate at a range that remains within desired tolerances for operating electric motors or other components of an aircraft. In addition, a battery pack may advantageously serve as an auxiliary source of power to drive motors or other components of an aircraft in the event of a fault in the generator(s) or other component of a hybrid-electric genset. This may therefore add a level of system safety and fault tolerance.
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 one or more battery packs and/or supercapacitors may be included as part of the hybrid-electric genset 161 and connected directly to the DC bus within the hybrid-electric genset 161, whether or not the aircraft components have separate batteries and/or supercapacitors. While
In various embodiments, fewer, additional, or different elements to those shown in
At an operation 213, a controller may therefore be able to determine how a power output of a hybrid-electric genset should be adjusted to maintain a desired voltage range on a DC bus. For example, if a battery's charge level is too low such that it is in danger of not being able to maintain a desired voltage, the controller may transmit instructions at an operation 217 to increase the power output of the hybrid-electric genset so that there is sufficient power to charge the battery. In another example, if a motor of the aircraft is currently using or is expected to require significantly more power than is currently being used, the controller may transmit instructions at an operation 217 to increase power output of the hybrid-electric genset. The power output may also similarly be decreased. In either instance, the controller may adjust this overall power output to the DC bus by varying the RPM supplied to an electric generator by an engine. As such, while the battery packs and supercapacitors may reduce a need to provide real time adjustments to power output of a hybrid-electric genset, as the battery packs and/or supercapacitors may maintain the DC bus at a desired voltage level, some control or adjustment of the RPM and therefore output power to the DC bus may still be desirable in various embodiments.
At an operation 303, aspects of power available at or flowing through a DC bus is measured by the controller. If the DC bus is measurable by a system-wide aircraft controller, the operation 303 may be carried out by the system-wide aircraft controller as well. Similarly, if batteries and/or supercapacitors are packaged as part of a hybrid-electric genset rather than being positioned as part of an overall aircraft system, the controller may at operation 303 also measure a state of the batteries/supercapacitors (e.g., charge state, current, voltage, etc.). At an operation 307, the controller determines how power output of the hybrid-electric genset should be adjusted based on the measurements. For example, if a DC bus voltage is getting close to going outside of a desired range, it may be desirable to transmit instructions at an operation 309 to the components of the hybrid-electric genset to adjust power output of the hybrid-electric genset based on the determination at the operation 307 to ensure the DC bus voltage stays within a desired voltage range.
In its most basic configuration, 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, memory 104 may be volatile (such as RAM 110), non-volatile (such as ROM 108, flash memory, etc.) or some combination of the two. Computing system environment 100 may have additional features and/or functionality. For example, computing system environment 100 may also include additional storage (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 by means of, for example, a hard disk drive interface 112, a magnetic disk drive interface 114, and/or an optical disk drive interface 116. As will be understood, these devices, which would be linked to the system bus 106, respectively, allow for reading from and writing to a hard disk 118, reading from or writing to a removable magnetic disk 120, and/or for reading from or writing to a removable optical disk 122, such as a CD/DVD ROM or other optical media. The drive interfaces and their associated computer-readable media allow for the nonvolatile storage of computer readable instructions, data structures, program modules and other data for the computing system environment 100. Those skilled in the art will further appreciate that other types of computer readable media that can store data may be used for this same purpose. Examples of such media devices include, but are not limited to, magnetic cassettes, flash memory cards, digital videodisks, Bernoulli cartridges, random access memories, nano-drives, memory sticks, other read/write and/or read-only memories and/or any other method or technology for storage of information such as computer readable instructions, data structures, program modules or other data. Any such computer storage media may be part of computing system environment 100.
A number of program modules may be stored in one or more of the 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, RAM 110, hard drive 118, and/or peripheral memory devices may be used to store computer executable instructions comprising an operating system 126, one or more applications programs 128 (which may include the functionality disclosed herein, for example), other program modules 130, and/or program data 132. Still further, computer-executable instructions may be downloaded to the computing environment 100 as needed, for example, via a network connection.
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. While not illustrated, other input devices may include a microphone, a joystick, a game pad, a scanner, etc. These and other input devices would typically be connected to the processing unit 102 by means of a peripheral interface 138 which, in turn, would be coupled to bus 106. Input devices may be directly or indirectly connected to processor 102 via interfaces such as, for example, a parallel port, game port, firewire, or a universal serial bus (USB). To view information from the computing system environment 100, a monitor 140 or other type of display device may also be connected to bus 106 via an interface, such as via video adapter 142. In addition to the monitor 140, the computing system environment 100 may also include other peripheral output devices, not shown, such as speakers and printers.
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 the remote computing system environment may be exchanged via a further processing device, such a network router 152, that is responsible for network routing. Communications with the network router 152 may be performed via a network interface component 154. Thus, within such a networked environment, e.g., the Internet, World Wide Web, LAN, or other like type of wired or wireless network, it will be appreciated that program modules depicted relative to the computing system environment 100, or portions thereof, may be stored in the memory storage device(s) of the computing system environment 100.
The computing system environment 100 may also include localization hardware 186 for determining a location of the computing system environment 100. In some instances, the localization hardware 156 may include, for example only, 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.
While this disclosure has described certain embodiments, it will be understood that the claims are not intended to be limited to these embodiments except as explicitly recited in the claims. On the contrary, the instant disclosure is intended to cover alternatives, modifications and equivalents, which may be included within the spirit and scope of the 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 obvious to one of ordinary skill in the art that systems and methods consistent with this disclosure may be practiced without these specific details. In other instances, well known methods, procedures, components, and circuits have not been described in detail as not to unnecessarily obscure various aspects of the present disclosure.
Some portions of the detailed descriptions of this disclosure have been presented in terms of procedures, logic blocks, processing, 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, etc., is herein, and generally, conceived to be a self-consistent sequence of steps or instructions leading to a desired result. The 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 is referred to as bits, values, elements, symbols, characters, terms, numbers, or the like, with reference to various presently disclosed embodiments.
It should be borne in mind, however, that these terms are to be interpreted as referencing physical manipulations and quantities and are merely convenient labels that should be interpreted further in view of terms commonly used in the art. Unless specifically stated otherwise, as apparent from the discussion herein, it is understood that throughout discussions of the present embodiment, discussions utilizing terms such as “determining” or “outputting” or “transmitting” or “recording” or “locating” or “storing” or “displaying” or “receiving” or “recognizing” or “utilizing” or “generating” or “providing” or “accessing” or “checking” or “notifying” or “delivering” or the like, refer to the action and processes of a computer system, or similar electronic computing device, that manipulates and transforms data. The data is represented as physical (electronic) quantities within the computer system's registers and memories and is transformed into other data similarly represented as physical quantities within the computer system memories or registers, or other such information storage, transmission, or display devices as described herein or otherwise understood to one of ordinary skill in the art.
Described herein are various embodiments for reducing noise emitted by an aircraft component such as an aircraft powerplant or component thereof. Although several embodiments described herein relate to enclosures for aircraft powerplants, such as engine cowlings, the various embodiments described herein may be used for components of aircraft other than powerplants and engines, and further still may be used to reduce noise emitted from components other than those of aircraft (e.g., helicopters, airplanes, vertical takeoff and landing (VTOL) aircraft, short takeoff and landing aircraft (STOL), etc.). For example, the embodiments described herein may also be implemented for any source of noise through which or around which air may pass, such as components of boats, motorcycles, automobiles, any other motor vehicle, or even for stationary components that generate noise around which or through which air passes.
Inlet and outlet airflow of noise producing components such as aircraft powerplants can be sources of noise, as the airflow into or out of those components can act as a medium through which noise and vibration can propagate. For example, in an aircraft having a hybrid powerplant, that hybrid powerplant may include a piston, rotary, or turbine engine that emits noise and has inlet and outlet airflow through which that noise may travel. Described herein are various embodiments for designing the geometry of the inlet and/or outlet airflow to reduce the amount of noise that is ultimately emitted from an enclosure having a noise emitting component therein, such as the cowling of an engine. For example, the various embodiments described herein, different air inlets and/or outlets may be configured to have a desired aspect ratio (e.g., length/width aspect ratio), eliminate line of sight from an engine (e.g., a noisy combustion engine) to any directions outside the aircraft that are noise sensitive, and/or line any internal noise-reflective surfaces with noise attenuating materials. Such embodiments as described herein advantageously provide for a weight-efficient and effective means of reducing operating noise from a noise emitting component, such as an aircraft hybrid powerplant.
The noise reducing embodiments described herein may be particularly advantageous for use in certain implementations. For example, some aircraft may have a hybrid powerplant that includes a combustion engine (e.g., turbine, rotary, piston) as well as an electric machine such as an electric motor/generator. Noise from such a hybrid generator will may be generated in and travel via the exhaust stream from the combustion engine, and further may escape via an airflow inlet for the combustion engine as well. While noise in an exhaust stream may be minimized with methods such as a muffler, other components of a hybrid powerplant and/or combustion engine may also generate noise, such as throughout the engine core, any cooling fans, pumps, and/or other accessory devices. Since that noise may generated at multiple places at once (e.g., from multiple sources/components), the noise may be hard to minimize.
Thus, the embodiments described herein are configured to reduce noise emitted by multiple sources (e.g., multiple powerplant or engine components that emit noise simultaneously). Noise may be carried from a noise source to a human ear via a medium such as air. The embodiments herein include enclosing noise sources (e.g., an aircraft powerplant) and manage airflow into and out of such an enclosure (and subsequently to the aircraft powerplant). The embodiments described herein further provide for additional noise reduction through the addition of noise attenuating material to various portions of the enclosure and in various configurations to reduce noise that may escape the enclosure (including noise that may escape through air inlets or outlets of the enclosure). That noise attenuating material may be a noise attenuating foam or any other type of suitable material.
Such noise attenuating material may further be placed within the enclosure (e.g., within the air inlet and/or outlet) in specific orientations and/or geometries to limit noise while also not hindering overall system performance (e.g., not hindering airflow to or from the powerplant). Accordingly, described herein are also orientations and geometries that are advantageously sized so as to not introduce unwanted pressure loss to an inlet or exit cooling airflow stream (e.g., backpressure to an exit cooling airflow stream). In various embodiments, noise attenuating materials used may also be selected based on their noise attenuating properties, resistance to fluids, heat, and/or fire, resistance to humidity, resistance to mold, resistance to corrosion, etc., so that the noise attenuating material has properties that are desirable for a given application.
Advantageously, the embodiments described herein therefore enable noise emitting components to be operated with a lower noise signature, which may be desirable, for example, in hybrid electric power for aviation.
As just one example of a system where it may be advantageous to use the systems and methods described herein, a hybrid powerplant configured to generate electrical and mechanical power for an aircraft may emit or produce noise that is desirable to minimize. For example, such a hybrid powerplant may include a prime mover such as an engine using combustion to create shaft work/power. That combustion may create noise, and noise from combustion engines in other applications is often released to an environment directly or conditioned using a muffler or similar method.
However, in addition to the noise created by the prime mover (e.g., combustion engine), a hybrid powerplant may have one or more other sources of noise, which may include, but are not limited to: (i) fuel injectors opening and closing, (ii) pistons slapping the cylinder walls inside a piston engine, (iii) fans whipping the air and/or slot gaps on fans creating noise, (iv) fluid pumps (e.g., oil, water, fuel), and/or (v) mechanical vibrations traveling through various parts and pieces.
The collective noise of the above and any other components of a hybrid powerplant may be referred to herein as the ambient noise of running a hybrid powerplant. Such ambient noise emission and propagation to a surrounding environment may be greatly reduced using the various systems and methods described herein.
A hybrid powerplant may have another feature that permits noise emitted from the powerplant to be released into an environment. An air intake or inlet for cool air may be used for combustion in the engine and/or other cooling tasks of the powerplant. An air exhaust or outlet for warm or hot air may also be released into the atmosphere. Since airflow velocities at intakes and exhaust for an engine are typically low compared to the speed of sound (e.g., less than 0.3 Mach (Ma)), any noise generated by components of the hybrid powerplant may travel through either intake and/or exhaust airstreams of the engine. In other words, the air moving into and out of a hybrid powerplant may carry sound waves, and the systems and methods described herein advantageously describe geometries and materials for air intake and/or exhausts that provide for significant reduction of noise emitted from an enclosure (e.g., a cowling) for a noise emitting component (e.g., a hybrid powerplant, combustion engine, related components of the combustion engine, etc.).
As further shown in
The sidewalls 147 of
The noise reduction chamber 173 may include noise attenuating elements, such as channels formed by vertically oriented walls within the noise reduction chamber. Examples of such channels are shown in and described further with respect to
Merely by way of example, various components 123, 129, 135, and 179, such as components of a hybrid powerplant for an aircraft, may be mounted or otherwise located at different positions within the cavity 117. More or less components may be typically included in the cavity 117, and various components may be in different locations with the cavity 117 than is shown in
Because different components 123, 129, 135, and 179 may be in different locations within the cavity 117, those components 123, 129, 135, and 179 may produce or emit noise that is emitted from different locations within the cavity. Therefore, as discussed herein, it may be difficult to specifically tune noise reduction elements for each and every potential source of noise within the enclosure 103. Thus, the noise reduction chamber 173 may attenuate noise or vibration propagating in the exhaust air as it travels through the noise reducing chamber 173 to the air outlet 157 (or may attenuate noise or vibration propagating in inlet air as it travels from the air inlet to the components in the cavity 117, in embodiments where noise attenuating elements (e.g., a noise reduction chamber) is placed along an air inlet path). For example, the plurality of channels in the noise reducing chamber may be formed of noise attenuating material, such that noise or vibration is absorbed by the walls of those channels, thereby reducing the amount of noise or vibration that is present in any air output at the air outlet 157.
As shown in
The noise reduction chamber 173 may also have a height D. A plurality of 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 space of the noise reduction chamber 173 (e.g., as shown in
As such, the plurality of walls within a noise reduction chamber, or other portion of an enclosure or cowling, may be arranged in any manner desired to achieve noise attenuation. The varying possible sizes of the plurality of walls and their associated channels may be referred to based on different aspect ratios applied to the geometry of the walls and channels. For example, a length/width ratio of a second section only (e.g., the part of the noise reducing chamber 208 that sticks out the back of the cavity) may be a length B of
These aspect ratios may be advantageously configured to create channels with desirable aspect ratios from a perspective of permitting adequate airflow through a noise reducing chamber. For example, low pressure drop passage of air may be desired either at an enclosure input or output. On one hand, if the channels formed by parallel walls are too wide (e.g., if the spacing of the foam compared to the length and height of the channel is too broad) then noise reduction qualities may be reduced. On the other hand, if the channels are narrow and very long, there is ample opportunity for the pressure waves of the noise to be attenuated by coming into contact with the plurality of walls. Accordingly, for a given application, wall material type, etc., a balance of channel width (e.g., length E or distance between two parallel planes), height (e.g., length D), and length (e.g., length A or distance along the axis of airflow principal direction) is important to advantageously achieve to balance desirable noise reduction qualities of the channels without meaningfully affecting performance of the engine or other components within the enclosure.
One example noise attenuating material that may be used in the embodiments described herein includes a melamine open-cell foam made from melamine resin. This foam may be characterized by excellent noise absorption with high fire retardancy and resistance to flame and smoke. For example, open-cell or closed-cell foams may be used, and may be formed from varying materials such as melamine, cellulose, polyethylene, cotton, any other suitable material, or any combination thereof. The plurality of walls described herein may have any desired thickness, and merely by way of example, thicknesses of one (1) inch to two (2) inches may be used. If different materials are used as the noise attenuating material, the thickness may be varied based on the properties of that material or combination of materials. The noise attenuating material and/or walls described herein may also be lined/coated with another material or may not be lined/coated with any other material. The walls configured to attenuate noise described herein may also be patterned in different ways to reduce resistance for airflow (e.g., smoother patterns) and/or increase noise attenuation. For example, the materials described herein may be formed to have a smooth surface, egg-crate surface pattern, pyramid-shaped surface pattern, wedge-shaped surface pattern, hemisphere-shaped surface pattern, wave-shaped surface pattern, any other pattern, or any combination thereof.
In the embodiment shown in
The walls are further arranged to create channels as described herein, such that air may pass between parallel or substantially parallel planes of foam. In this way, sound pressure waves may be attenuated while the core flow of air through a noise reduction chamber has minimal restriction as it heads into or out of the system. As such, it is desirable to configure the walls and the channels between them such that there is not too much aerodynamic resistance (e.g., pressure loss) airflow streams (e.g., for air flow used for cooling engine components), then performance of an engine (including e.g., performance of cooling systems) may degrade. With properly sized channels such an effect of decreased performance may be minimized. In various embodiments, as described herein, non-parallel walls may additionally or alternatively be used.
In particular,
In an illustrative 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 memory. Upon execution of the computer-readable instructions by a processor, the computer-readable instructions may cause a computing device to perform the operations.
The foregoing description of illustrative embodiments has been presented for purposes of illustration and of description. It is not intended to be exhaustive or limiting with respect to the precise form disclosed, and modifications and variations are 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 claims appended hereto and their equivalents.
This application is a continuation of PCT/US2022/017379, filed Feb. 22, 2022, which claims the benefit of U.S. Provisional Patent Application No. 63/280,615, filed Nov. 17, 2021, U.S. Provisional Patent Application No. 63/163,165, filed Mar. 19, 2021, and U.S. Provisional Patent Application No. 63/151,760, filed Feb. 21, 2021, the entire contents of each of which are hereby incorporated by reference in their entirety.
| Number | Date | Country | |
|---|---|---|---|
| 63280615 | Nov 2021 | US | |
| 63163165 | Mar 2021 | US | |
| 63151760 | Feb 2021 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | PCT/US22/17379 | Feb 2022 | US |
| Child | 18236363 | US |
