REVERSE FLOW GAS TURBINE ENGINE HAVING ELECTRIC MACHINE

An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and thereby form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.

PRIORITY INFORMATION

The present application claims priority to U.S. patent application Ser. No. 18/307,938, filed on Apr. 27, 2023, which claims priority to Polish Patent Application Number P.443814, filed on Feb. 17, 2023. U.S. patent application Ser. No. 18/307,938 and Polish Patent Application Number P.443814 are hereby incorporated by reference in their entirety for all purposes.

FIELD

The present disclosure relates generally to a reverse flow gas turbine engine having an electric machine.

BACKGROUND

A gas turbine engine generally includes a turbomachine and a rotor assembly. Gas turbine engines, such as turboprop engines, may be used for aircraft propulsion. In the case of a turboprop engine, the rotor assembly may be configured as a variable pitch propeller. In some installations the gas turbine engine is oriented in a reverse flow configuration such that an air flow provided by forward motion of an aircraft is received by the gas turbine engine and turned to flow in a reverse direction through the turbomachinery of the gas turbine engine before an exhaust is discharged from the engine. Locating auxiliary components in such a reverse flow configuration remains an area of interest.

DETAILED DESCRIPTION

The term “at least one of” in the context of, e.g., “at least one of A, B, and C” refers to only A, only B, only C, or any combination of A, B, and C.

The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source.

The term “combustion section” refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section including one or more of a deflagrative combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other appropriate heat addition assembly. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof.

The terms “low” and “high”, or their respective comparative degrees (e.g., -er, where applicable), when used with a compressor, a turbine, a shaft, or spool components, etc. each refer to relative pressure within an engine unless otherwise specified. For example, a “low turbine” or “low pressure turbine” defines a component configured to operate at a pressure lower than a “high pressure turbine” of the engine.

The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operational attitude of a vehicle such as an aircraft. For example, with regard to an aircraft, forward refers to a position closer to a nose of the aircraft and aft refers to a position closer to an empennage of the aircraft.

As used herein, “top” refers to a highest or uppermost point, portion, or surface of a component in the orientations shown in the figures.

As used herein, “bottom” refers to a lowest or lowermost point, portion, or surface of a component in the orientations shown in the figures.

As used herein, “remote” or a “remote” component means a component that is discrete from a first component such that the component is free from mechanical coupling (e.g., direct or indirect) to the first component. For example, a remote propulsor is a propulsor that is discrete from a first propulsor such that the remote propulsor can be mounted at a location on the aircraft that is different than the location of the first propulsor. The remote component may be communicatively or electrically coupled to the first component.

As used herein, a “propulsor” is a component that rotates and generates thrust. The propulsor can be a propeller. For example, the propulsor can be a propeller of a turboprop engine that is drivingly coupled to the turbo-engine such that rotation of the components of the turbo-engine causes the propulsor to rotate and to generate thrust. The propulsor can be an electric propulsor (e.g., propeller) that is remote from the turbine engine and is drivingly coupled to an electric machine such that the electric machine causes the propulsor to rotate and to generate thrust.

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about,” “approximately,” “generally,” and “substantially” is not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or the machines for constructing the components and/or the systems or manufacturing the components and/or the systems. For example, the approximating language may refer to being within a one, two, four, ten, fifteen, or twenty percent margin in either individual values, range(s) of values and/or endpoints defining range(s) of values.

As will be discussed in more detail below, the subject matter of the present disclosure is directed generally to locating an electric machine near an intake channel of a reverse flow turboprop engine at a location which is closer to an aft end of the turboprop engine than to a forward end of the turboprop engine. The electric machine is rotatingly coupled to a low pressure shaft of the reverse flow turboprop engine and as a consequence of the reverse flow configuration the low pressure shaft extends aft of a core of the turboprop engine. The electric machine can be operated as a generator and/or motor for use in either adding power to and/or extracting power from the low pressure shaft. Placement of the electric machine in the proximate location described above permits an exchange of heat between the electric machine and a flow of air traversing through an intake channel of the engine. Such exchange of heat can provide tighter packaging of the electric machine and/or higher heat generating operating demands placed upon the electric machine.

To accommodate the placement of the electric machine in an aft location in at least one embodiment, an intake channel may be provided that forms a non-annular flow path at an inlet to the intake channel which then changes to an annular flow path around the LP shaft prior to air being delivered to a compressor of the gas turbine engine.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,FIG.1is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment ofFIG.1, the gas turbine engine is a reverse flow turboprop engine10, referred to herein as “turboprop engine10.” As shown inFIG.1, turboprop engine10defines an axial direction A (extending parallel to a longitudinal centerline or central axis12provided for reference), a radial direction R, and a circumferential direction C (not shown) disposed about the axial direction A. Turboprop engine10generally includes a propeller section14and a core turbine engine16disposed aft of the propeller section14from an aircraft perspective, the propeller section14being operable with, and driven by, core turbine engine16. The exemplary core turbine engine16depicted generally includes a substantially tubular outer casing18extending generally along axial direction A. Outer casing18generally encloses core turbine engine16and may be formed from a single casing or multiple casings. Core turbine engine16includes, in a serial flow relationship, a compressor22, a combustion section26, a high pressure (HP) turbine28, a low pressure (LP) turbine30, and an exhaust section32. An air flow path generally extends through compressor22, combustion section26, HP turbine28, LP turbine30, and exhaust section32which are in fluid communication with each other.

An HP shaft or spool34drivingly connects the HP turbine28to the compressor22. An LP shaft or spool36drivingly connects the LP turbine30to propeller section14of the turboprop engine10. For the embodiment depicted, propeller section14includes a variable pitch propeller38having a plurality of propeller blades40coupled to a disk42in a spaced apart manner. As depicted, the propeller blades40extend outwardly from disk42generally along the radial direction R. Each propeller blade40is rotatable relative to the disk42about a pitch axis P by virtue of the propeller blades40being operatively coupled to a suitable actuation member44configured to collectively vary the pitch of the propeller blades40in unison. The propeller blades40, disk42, and actuation member44are together rotatable about the longitudinal centerline12by LP shaft36across a power gear box46. The power gear box46includes a plurality of gears for stepping down the rotational speed of the LP shaft36to a more efficient rotational fan speed and is attached to one or both of a core frame or a fan frame through one or more coupling systems. Disk42is covered by a rotatable front hub48aerodynamically contoured to promote an airflow through the plurality of propeller blades40.

During operation of the turboprop engine10, a volume of air50(also referred to as a free stream flow of air51prior to its encounter with the propeller38, and referred to as an incoming flow of air50after passage through the propeller38) passes through blades40of propeller38and is urged toward a radial inlet52of core turbine engine16. More specifically, turboprop engine10includes an intake channel54that defines radial inlet52that routes an inlet portion of air53of the flow of air50from inlet52downstream to compressor22. Though the inlet52is depicted as a radial inlet in the embodiment ofFIG.1, other configurations of inlet52are also contemplated. For example, the inlet52can also take the form of an inlet arranged in an axial direction to capture the inlet portion of air53of the volume of air50. The inlet portion of air53of the flow of air50captured by the inlet52is referred to herein as an intake flow of air. The intake channel54defines the intake flow of air and generally extends from an inlet of the intake channel54to just upstream of the compressor22.

The turboprop engine10embodiments described herein are configured as reverse flow engines. Such engines are characterized by a general relationship between the direction of the flow of incoming air50(such direction can be used to characterize the relative motion of air during a mode of operation of the engine10such as a forward thrust mode) and that of the flow of air axially through the turboprop engine10. The flow of air through the core turbine engine16is generally reverse to that of the flow of incoming air50. Turning the flow from the direction of the incoming flow of air50to the axial direction through the core turbine engine16is usually performed by the intake channel54. The change of direction is reversed in that the bulk direction of the flow of air50(itself having a circumferential swirl component imparted by the propeller blades40in addition to a longitudinal component) is opposite, or reverse, to the bulk direction of air flow axially through the core turbine engine16(which itself also includes a longitudinal component but also include radial and circumferential components owing to the shape of the flow path and swirl induced by rotating turbomachinery components) during one or more phases of operation of the core turbine engine16. Thus, it will also be appreciated that the term “reverse” is a relative comparison of the longitudinal components of the bulk flow of air50and bulk flow of air axially within the engine10. Though the longitudinal direction of the flow of air50may not be perfectly parallel with the axial flow of air through the engine10, it will be appreciated that the longitudinal components of the directions the flow of air50and the axial flow are reversed.

Compressor22includes one or more sequential stages of compressor stator vanes60, one or more sequential stages of compressor rotor blades62, and an impeller64. Though the illustrated embodiment includes both axial and centrifugal flow compressors, in some forms the turboprop engine10can include just an axial flow compressor(s) or centrifugal flow compressor(s). The one or more sequential stages of compressor stator vanes60are coupled to the outer casing18and compressor rotor blades62are coupled to HP shaft34to progressively compress the air53. Impeller64further compresses air53and directs the compressed air53into combustion section26where air53mixes with fuel. Combustion section26includes a combustor66which combusts the air/fuel mixture to provide combustion gases68.

Combustion gases68flow through HP turbine28which includes one or more sequential stages of turbine stator vanes70and one or more sequential stages of turbine blades72. The one or more sequential stages of turbine stator vanes70are coupled to the outer casing18and turbine blades72are coupled to HP shaft34extract thermal and/or kinetic energy therefrom. Combustion gases68subsequently flow through LP turbine30, where an additional amount of energy is extracted through additional stages of turbine stator vanes70and turbine blades72coupled to LP shaft36. The energy extraction from HP turbine28supports operation of compressor22through HP shaft34and the energy extraction from LP turbine30supports operation of propeller section14through LP shaft36. Combustion gases68exit turboprop engine10through exhaust section32.

It will be understood that one or more rows of stator vanes60and70can be variable vanes controlled by a controller (see below with respect to controller100) in one form. Furthermore, with particular respect to stator vanes70, one or more rows of the stator vanes70can be variable.

In other exemplary embodiments, the turbine engine may include any suitable number of compressors, turbines, shafts, etc. For example, as will be appreciated, HP shaft34and LP shaft36may further be coupled to any suitable device for any suitable purpose. For example, in certain exemplary embodiments, turboprop engine10ofFIG.1may be utilized in aeroderivative applications. Additionally, in other exemplary embodiments, turboprop engine10may include any other suitable type of combustor, and may not include the exemplary reverse flow combustor depicted.

The embodiment of turboprop engine10illustrated inFIG.1includes an electric machine74located aft of the core turbine engine16and rotatingly coupled to the LP shaft36. In some forms the electric machine74is contained in an environmentally sealed housing which can be pressurized to minimize electrical corona and discharge effects. Further, in some forms the LP shaft36and electric machine74can be configured to rotate at a constant speed from idle to max power, with thrust of the engine10controlled by the variable pitch propeller38.

Given the coaxial relationship of the LP shaft36with the HP shaft34, in such an embodiment the LP shaft36is configured to extend aft of the core turbine engine16, and also further aft than the HP shaft34, despite the HP compressor22being the upstream-most compressor of the turboprop engine10(i.e., despite there being no low-pressure compressor upstream of the HP compressor22and downstream of the inlet52).

The electric machine74can be used in many different power configurations. In one form the electric machine is configured to extract power from the LP shaft36when the machine74operates as a generator. The extraction of mechanical power from the LP shaft36and conversion to electric power can be used to charge an on-board power storage device such as a battery, or alternatively to provide power to another electrical device (e.g., an electric motor, an electrical accessory on an aircraft, etc.). In other forms, the electric machine74can be used as a motor to provide power to the LP shaft36to supplement power extracted by the LP turbine30from the combustion gases68. In these forms, the electric machine74can be configured to provide a minimum of 10% of supplemental thrust to the engine10, a minimum of 20% of supplemental thrust to the engine10, and up to 40% of supplemental thrust to the engine10in various embodiments. In still other forms, the electric machine74can be configured to power to drive 100% of thrust from the propeller section14. A scenario in which the electric machine74provides all power to the propeller section14can include shutdown of the engine10. In one non-limiting example of an engine being shut down, upon or near landing the engine10can be command to shut down and the electric machine74used to drive further propeller thrust requirements, whether that includes fine power on short final or power when the propeller section14is configured in reverse pitch to aid in slowing the aircraft.

A battery or other secondary power source can be used to provide power to the electric machine74when operated as a motor. The supplementation of power by the electric machine74to the LP shaft36in this manner can be transitory or steady state, depending on the control requirements requested of the turboprop engine. For example, in those operating conditions in which power output of the engine lags behind a commanded power, the electric machine74can provide near instantaneous supplemental power to the LP shaft36to provide on-condition power output from the turboprop until the engine10achieves a steady state operating condition at the higher output power. In still further forms, the electric machine can be operated as a motor in some portions of operation of the engine10, and as a generator in other portions of operation of the engine10, along the lines of any of the variations discussed herein.

Given the proximity of the electric machine74to the intake channel54, the electric machine74can be further positioned to exchange heat with the inlet portion of the air53traversing the intake channel54to aid in removing beat from the electric machine74. Cooling of the electric machine74using the inlet portion of air53permits tighter packaging of the electric machine74and closer location of the electric machine74to heat generating portions of the engine10. The relative location of the electric machine74and intake channel54can permit an effective exchange of heat through any number of useful mechanisms including at least one of conduction and convection cooling. In some forms cooling air may directly impinge upon the electric machine74and/or may be used to vent a cavity in which the electric machine74is located. In still further forms the electric machine74can be used in part to form the flow path of the intake channel54to provide direct heat transfer between the electric machine74and the inlet portion of air53. In yet still further forms the electric machine74may be in direct contact with a portion of the intake channel54forming the flow path such that heat transfer occurs between the electric machine74and inlet portion of air53via that particular portion of the intake channel54forming the flow path. Further aspects of the location of electric machine74and various cooling techniques are described further below.

Various other embodiments are disclosed further herein related to the location and use of the electric machine74, the shape and configuration of the intake channel54, and various techniques to exchange heat between the electric machine74and the inlet portion of air53flowing in the intake channel54. As will be appreciated, like reference numerals refer to like elements and, thus, any of the variations disclosed herein related to any particular exemplary embodiment in any given figure are also applicable to embodiments depicted and discussed with respect to the other figures.

Turning now toFIG.2, an exemplary embodiment of the turboprop engine10is illustrated in which the engine10includes an electric machine74located aft of the core turbine engine16in a tail cone76defining an aft end of an engine nacelle78. As will be appreciated, the engine nacelle78is used to enclose the gas turbine engine and includes one or more portions that interface with aircraft structure such as a wing, pylon, fuselage, etc. The tail cone76may be a complete body of revolution that circumferentially encloses the electric machine74in some embodiments. In alternative embodiments, however, the tail cone76may be a partial body of revolution or other shape that covers the electric machine to complete an enclosure with other aircraft structure (e.g., wing, pylon, fuselage, etc.). Thus, the tail cone76is any suitable structure of the engine nacelle which is located aft of the core turbine engine16and is used to wholly or partially enclose the electric machine74.

In some forms of the embodiments disclosed herein the engine nacelle78may take on the form of an engine cowling when the engine10is installed on a single engine turboprop aircraft. In such installations the electric machine74can be located between the core turbine engine16and a firewall of the aircraft. In such installations, therefore, the engine cowling or other forebody structure of the aircraft can be used to enclose the electric machine74.

The embodiment ofFIG.2includes an accessory gear box (AGB)80located aft of the HP compressor22and is coupled to a starter motor82. The starter motor82is coupled to the HP shaft via the AGB80such that during a start sequence of the turboprop engine10the starter motor82can be used to impart rotational power via the AGB80to the HP shaft34. In the illustrated embodiment, the AGB80is depicted as being co-axial with the HP shaft34and LP shaft36(it will be appreciated that although the AGB80is rotatingly coupled to the HP shaft34, it is not otherwise rotatingly coupled with the LP shaft36). To provide such a coaxial relationship between the HP shaft34and AGB80, in one form the AGB80is a planetary gear system in which the HP shaft34is coupled to a sun gear of the planetary gear system. In other forms a central gear of the AGB80is coupled via one or more idler gears to the starter motor82. Other forms are also contemplated to permit a co-axial relationship between the AGB80and HP shaft34. In other forms, however, the AGB80need not be co-axial with the HP shaft34.

The electric machine74is also depicted inFIG.2as being coupled to the LP shaft36through a speed change device84which can be used to alter a speed ratio between the LP shaft36and the electric machine74. In some forms, the speed change device84can be a transmission that provides a fixed speed ratio, but in other forms the transmission can provide variable speed ratios. In still other forms, the transmission can include a clutch mechanism to disengage the electric machine74from the LP shaft36. Still further, although the embodiment depicted inFIG.2includes the speed change device84, other embodiments need not include the device84. In such an embodiment, the electric machine74is directly connected to the LP shaft36.

The intake channel54forms a flow path between the AGB80and the HP compressor22. The flow path of the intake channel54follows the route depicted inFIG.2, where the inlet portion of air53captured by the intake channel54passes first through a non-annular inlet90depicted at sightline A-A (an embodiment of which is illustrated inFIG.7A, discussed in more detail below) before it is split into an annular flow path92depicted at sightline B-B (an embodiment of which is illustrated inFIG.7B, discussed in more detail below). The non-annular inlet90is radially offset from the LP shaft36and confined to a circumferential section about the LP shaft36. The intake channel54takes the form of the annular shape leading up to the delivery of the inlet portion of air53to the HP compressor22. The intake channel54thus begins with a non-annular intake shape and ends with an annular shape. As will be appreciated, the annular shape of the intake channel includes a central interior that accommodates the LP shaft36. The intake channel54in the illustrated embodiment is thus required to morph, or change, from the non-annular shape at the inlet90to the annular shape at its discharge to the compressor22so that the intake channel54accommodates the intrusion of the LP shaft36through the intake channel54and to the electric machine74. In this way, the LP shaft36extends through at least a portion of the intake channel54where the flow path in the intake channel54changes shape to accommodate the LP shaft36. In some forms, an exterior surface of the LP shaft36may be exposed to the inlet portion of air53such that the inlet portion of air53flowing in the intake channel54is exposed to the rotating exterior surface of the LP shaft36. In other forms, however, the intake channel54may include a separate structure used to define the central interior and which is used to separate the exterior surface of the LP shaft36from the inlet portion of air53flowing in the intake channel54.

In some forms, the inlet90of the intake channel54is located at either a six o-clock position on the nacelle such as what would conventionally be considered the bottom, or underside, of the nacelle. An inlet on a Beechcraft Denali or Beechcraft King Air Turboprop are examples. The exhaust section32can be located at either or both of the three o'clock and nine o'clock position on the nacelle such as would conventionally be considered a left or right side of the nacelle. In this manner, the spacing of the inlet90of the intake channel54is circumferentially displaced from the exhaust section32to minimize/prevent exhaust gases from being circulated to the inlet90for ingestion into the engine10. Furthermore, it will be appreciated that the inlet52, though illustrated at an axially aft location in the various embodiments, can be located forward closer to the blades40while still maintaining the configuration to reverse the flow from the direction of the incoming flow of air50to the axial flow direction required in the turbine engine configurations depicted.

Also depicted inFIG.2is an offtake flow path86created by an offtake opening88provided in the intake channel54and which is configured to provide a flow of offtake air94to be used for heat exchange purposes with the electric machine74. The offtake opening88can be located downstream of the inlet90to the intake channel54and is structured to remove part of the inlet portion of air53flowing through the offtake channel86. The offtake opening88can be a permanent vent structure that includes a fixed opening through which air can pass regardless of mode of operation of the electric machine74. In other forms, however, the offtake opening88can include a movable mechanical structure that permits modulating the area of the offtake opening88, including in some forms fully closing the offtake opening. Such a movable mechanical structure can take any variety of forms such as a hinged plate, sleeve valve, or other suitable device.

The offtake flow path86can take a variety of forms including the solid line depicted inFIG.2in which the flow path86flows past the electric machine74, In one form, the offtake flow path86can alternatively and/or additionally be routed along the dotted line shown inFIG.2. Such a flow path can provide additional level of venting of the cavity in which the electric machine74is located, and/or provide greater dwell time within the cavity to ensure a higher level of heat exchange.

The offtake flow path can include one or more discharge openings96and98to permit the flow of offtake air94to exit from the nacelle78. The flow of offtake air94can be urged to exit through a pressure differential that exists between the offtake opening88and the discharge openings96and/or98. Such a pressure differential can be provide via ejector action if needed through suitable structure configured to provide such an action (e.g., a venturi ejector). One or both of the discharge openings96and98, in some embodiments, can be a permanent vent structure that includes a fixed opening through which air can pass regardless of mode of operation of the electric machine74.

In other forms, however, one or both of the discharge openings96and98can include a movable mechanical structure that permits modulating the area of the discharge openings96and98, including in some forms fully closing the discharge openings. Such a movable mechanical structure can take any variety of forms such as a hinged plate, sleeve valve, or other suitable device. The moveable mechanical structure can protrude into a passing flow of air to which the discharged flow of offtake air94is being discharged, and in other forms can protrude into the offtake flow path86.

Given the proximity of the electric machine74to the intake channel54and the configurations disclosed herein, various cooling techniques are contemplated with respect to the various embodiments. For example, though the passing flow of offtake air94is illustrated inFIG.2passing adjacent and/or around the electric machine74, in some embodiments, the passing flow of offtake air94can be directed to impinge directly upon a portion of the electric machine. Further, a surface of the electric machine74, such as an outer housing, can form part of the flow path of the offtake flow path86. In other forms, a structure forming the offtake flow path86can be in heat conductive relationship with a portion of the electric machine74(e.g., a housing of the electric machine74).

The embodiment depicted inFIG.2also includes a controller100configured to control various aspects of the depicted embodiment (the embodiment depicted inFIG.1can also include a controller for control of analogous features). As depicted through the various dotted lines, the controller100can control one or more different systems associated with operation of the engine10. The dotted nature depicted in the figure denotes the optional inclusion of one or more, or all, of the systems connected with the controller100. For example, the controller100can be used to control variable stator vanes70in either or both turbines28and30. Additionally and/or alternatively, the controller100can be used to control variable stator vanes60in the compressor22. Additionally and/or alternatively, the controller100can be used to control fuel flow to the combustion section26. Additionally and/or alternatively, the controller100can be used to control one or more of the openings of the openings88,96, and98. Additionally and/or alternatively, the controller100can be used to control operation of the electric machine74. Additionally and/or alternatively, the controller100can command fuel flow to the combustion section26to be stopped and also simultaneously command the propeller section14to be positioned in a forward or reverse pitch configuration.

The controller100can thus be used in any or all of the following examples. The controller100may selectively drive the electric machine74as a generator. In those situations in which the electric machine74is ‘powered on’ to operate as a generator, the controller100can make adjustments to fuel rate delivery to the combustion section26while also optionally changing position of variable stator vanes60and/or70. The controller100can optionally operate a clutch in the speed change device84.

Turning now toFIG.3, another embodiment of the reverse flow engine10discussed above is illustrated. The exemplary engine10ofFIG.3may be configured in substantially the same manner as the exemplary engine10ofFIG.2, and as such the same or similar numbers may refer to the same or similar parts.

For the embodiment ofFIG.3, the engine10includes an electric machine74coupled to the engine10and operated as a generator, the electric machine74further in electrical communication via a power conduit102with an electric machine74bwhich is operated as a motor. The electric machine74boperated as a motor is used to drive a set of propeller blades40bapart from the blades40adriven by the turbine engine10depicted at the top ofFIG.3. The electric machine74bconfigured as a motor can be used to provide additional thrust output beyond that provided by the propeller blades40adriven by the turbine engine10.

Although the electric machine74bis illustrated apart from any turbomachinery components such as those at the top of the figure, in some embodiments the electric machine74bcan be integrated with a gas turbine engine much in the same manner as the turboprop engine10depicted at the top of the figure. In these embodiments, the electrical coupling between the electric machines74aand74bcan be used to exchange power between the two (e.g., where one machine is a motor and the other a generator) or can be coupled to a common energy storage device (e.g. a bank of batteries). Any of the variations in the embodiments discussed above are also applicable to the embodiments shown inFIG.3, such as but not limited to heat exchange between the electric machine74and the inlet portion of air53, the offtake flow path86, etc.

Turning now toFIG.4, another embodiment of the reverse flow engine10discussed above is illustrated. The exemplary engine10ofFIG.4may be configured in substantially the same manner as the exemplary engine10ofFIG.2, and, as such, the same or similar numbers may refer to the same or similar parts.

For the embodiment ofFIG.4, the engine10includes an electric machine74coupled to the HP shaft34of the engine10, in which the configuration can employ the electric machine74as either a motor (e.g., to start the engine10) or a generator (e.g., to scavenge power for electric power generation). As will be appreciated in this embodiment, the LP shaft36need not be extended to the rear of the engine10.

Turning now toFIG.5, another embodiment of the reverse flow engine10discussed above is illustrated. The exemplary engine10ofFIG.5may be configured in substantially the same manner as the exemplary engine10ofFIG.2, and as such the same or similar numbers may refer to the same or similar parts.

For the embodiment ofFIG.5, the engine10includes electric machines74aand74b, each coupled with respective LP shaft36and HP shaft34. This embodiment enables power input/extraction to/from either spool34,36independently, as well as the potential to transfer power between the two spools34,36of the engine to improve operability/mitigate vibration, compressor stall or instability issues. As will therefore be appreciated, power can be extracted from spool34and provided to spool36in one mode of operation, power can be extracted from spool36and provided to spool34in another mode of operation, and power can be either extracted from or provided to both spools34,36in yet another mode of operation.

Any of the electric machines discussed herein are capable of being packaged so as to provide power densities suitable for use in the applications discussed herein. For example, the electric machines can have a power density ranging anywhere from greater than 3 kW/kg, greater than 5 KW/kg, and greater than 6 kW/kg.

As noted, the exemplary controller100useful in any ofFIGS.2-5is configured to regulate any of the aforementioned systems such as variable stator vanes60and/or70, electric machine74, fuel flow to the combustion section26, etc., either alone or in combination, based on a control scheme stored in the controller100. In one or more exemplary embodiments, the controller100depicted inFIGS.2and3may be a stand-alone controller100for any of the aforementioned systems, or alternatively, may be integrated into one or more of a controller for the gas turbine engine with which the aforementioned systems are integrated, a controller for an aircraft including the gas turbine engine with which the aforementioned systems are integrated, etc.

Referring particularly to the operation of the controller100, in at least certain embodiments, the controller100can include one or more computing device(s)104such as depicted inFIG.6. The computing device(s)104can include one or more processor(s)104A and one or more memory device(s)104B. The one or more processor(s)104A can include any suitable processing device, such as a microprocessor, microcontroller, integrated circuit, logic device, and/or other suitable processing device. The one or more memory device(s)104B can include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, RAM, ROM, hard drives, flash drives, and/or other memory devices.

The one or more memory device(s)104B can store information accessible by the one or more processor(s)104A, including computer-readable instructions104C that can be executed by the one or more processor(s)104A. The instructions104C can be any set of instructions that when executed by the one or more processor(s)104A, cause the one or more processor(s)104A to perform operations. In some embodiments, the instructions104C can be executed by the one or more processor(s)104A to cause the one or more processor(s)104A to perform operations, such as any of the operations and functions for which the controller100and/or the computing device(s)104are configured, the operations for any of the aforementioned systems such as variable stator vanes60and/or70, electric machine74, fuel flow to the combustion section26, etc., as described herein, and/or any other operations or functions of the one or more computing device(s)104. The instructions104C can be software written in any suitable programming language or can be implemented in hardware. Additionally, and/or alternatively, the instructions104C can be executed in logically and/or virtually separate threads on the one or more processor(s)104A. The one or more memory device(s)104B can further store data104D that can be accessed by the one or more processor(s)104A. For example, the data104D can include data indicative of power flows, data indicative of engine/aircraft operating conditions, and/or any other data and/or information described herein.

The computing device(s)104can also include a network interface104E used to communicate, for example, with the other components of system (e.g., via a communication network). The network interface104E can include any suitable components for interfacing with one or more network(s), including, for example, transmitters, receivers, ports, controllers, antennas, and/or other suitable components. One or more devices can be configured to receive one or more commands from the computing device(s)104or provide one or more commands to the computing device(s)104.

The network interface104E can include any suitable components for interfacing with one or more network(s), including, for example, transmitters, receivers, ports, controllers, antennas, and/or other suitable components.

Turning now toFIGS.7A and7B, cross sectional views of two separate flow stations along the flow path of the intake channel54are illustrated.FIG.7Aillustrates view A-A fromFIGS.2and3, above, which depicts a cross section of the intake channel54that routes an inlet portion of air53of the flow of air50. The cross-section A-A depicted inFIG.7Ais at or close to the inlet of the intake channel54and has a non-annular flow shape. The non-annular flow shape depicted inFIG.7Ais similar to a kidney shape in the illustrated embodiment, but other non-annular shapes are also contemplated. The cross-section B-B depicted inFIG.7B, fromFIGS.2and3above, is at or close to the discharge of the intake channel54as it delivers the inlet portion of air53of the flow of air50to the compressor22. The annular flow shape depicted inFIG.7Bencloses the LP shaft36.

Referring now toFIG.8, a flow diagram of a method of operating a gas turbine engine in accordance with an exemplary aspect of the present disclosure is provided. The method ofFIG.7may be utilized to operate one or more of the exemplary engine and electric machine placement described above with reference toFIGS.1through5B. Accordingly, it will be appreciated that the method may generally be utilized to operate a gas turbine engine including engine10, electric machine74, and intake channel54. However, in other exemplary aspects, the method may additionally or alternatively be utilized to operate any other suitable gas turbine engine.

More specifically,FIG.8discloses a method106of operating a reverse flow gas turbine engine which includes at108operating a gas turbine engine having a reverse flow configuration. The engine can take the form of a turboprop as discussed above in various embodiments. Step110includes receiving an incoming flow of air into an intake channel of the gas turbine engine. The flow of air received in the intake channel is turned at step112from its initial flow direction into an axial flow direction of the gas turbine engine. Step114discloses cooling an electric machine located aft of a turbine of the gas turbine engine using the flow of air in the intake channel.

The arrangement of the electric machine74coupled with the low pressure shaft36and located on a side of the high pressure compressor22opposite the high pressure turbine28provides various technical effects, including the ability to cool the electric machine74using the intake flow of air. Such a placement provides for additional separation from hot section components of the engine10. The electric machine74can be placed in proximity to the intake channel54at an aft end of the engine10in a heat exchange relationship, where the intake flow of air can exchange heat through either or both of conduction and convection with the electric machine74. In some forms an offtake flow of air can be extracted from the intake flow of air for use in cooling the electric machine74in lieu of the intake flow of air for additional flexibility, such as selective cooling provided through activation of discharge openings96,98. Various other flexible arrangements can also be provided of the electric machine74, and specifically cooling of the electric machine. For example, an electric machine74bcan be coupled with the low pressure shaft36, while another electric machine74ais coupled with the high pressure shaft34. Both of electric machines74aand74bcan be located on a side of the high pressure compressor22opposite the high pressure turbine28.

As noted above, turbine engines include a turbo-engine that provides power for rotating a propulsor (e.g., a propeller). The turbo-engine burns air and fuel to generate an exhaust gas flow that drives an aerodynamically-coupled power turbine, or low-pressure turbine, that is further coupled to a set of propulsor blades of the propulsor via a propulsor shaft. In this sense, a speed of the turbo-engine indirectly affects a speed of the power turbine by providing torque to drive the propulsor, thus providing thrust.

The pitch of the propulsor blades can also be rotated on the propulsor shaft to provide additional thrust at the expense of increasing the torque demand of the turbine shaft. For instance, if the pitch of the propulsor blades is reduced, each propulsor blade rotates on the propulsor shaft such that the air resistance of the propulsor blade as the propulsor blade rotates about the propulsor shaft decreases, and, thus, reduces the torque demand for the propulsor shaft. In the instance when the turbo-engine speed is constant, the reduced torque demand of the propulsor results in increased propulsor shaft rotational speed (RPM). Conversely, if the pitch of the propulsor blades is increased, each propulsor blade rotates on the propulsor shaft such that the air resistance of the propulsor blade as the propulsor blade rotates about the propulsor shaft increases, and, thus, increases the torque demand for the propulsor shaft. In the instance when the turbo-engine speed is constant, the increased torque demand of the propulsor results in decreased propulsor shaft speed.

In certain embodiments, the present disclosure provides for a propulsion system including a turbine engine (e.g., a turboprop engine) with a propulsor, and a remote propulsor that is remote from the turbine engine. The turbine engine includes an electric machine and the remote propulsor includes an electric machine. In this way, the propulsor of the turbine engine can be powered electrically by the electric machine. Similarly, the remote propulsor can be powered electrically by the electric machine. In some embodiments, the remote propulsor is coupled to a remote turbine engine such that the remote turbine engine powers the remote propulsor.

Turbine engine control systems typically employ dual throttle levers used by the pilot to adjust, respectively, the turbo-engine speed demand, e.g., the speed of the turbo-engine, and the propulsor speed demand, e.g., via the pitch of the propulsor blades. In some instances, the dual throttle levers include a first throttle lever that controls the turbine engine and a second throttle lever that controls the remote propulsor. This dual-lever turbine engine control system allows for variation of aircraft speed and torque demand during, for instance, takeoff, cruise, or reverse thrust operations. Typically, small turboprop aircraft include one lever for the core engine speed (power or fuel flow), and a second lever for controlling the pitch of the propulsor blades. The propulsor is designed to operate at a constant speed, so a pilot would need to first adjust the core engine speed to increase/decrease power, and then adjust the propulsor blade pitch to maintain the proper propulsor speed. The propulsor is driven by a free turbine downstream of the core engine, so when the core speed increases, if the propulsor pitch remains constant, then the propulsor speed increases. The pilot needs to change the pitch to counter the change in speed resulting from the core engine change. This results in an increased workload for the pilot.

The present disclosure provides for an improved turbine engine control system that receives a single throttle lever input, and controls the turbine engine speed and the propulsor speed and the propulsor pitch angle based on the single throttle lever input. The turbine engine control system can also control the remote propulsor speed and the remote propulsor pitch angle based on the single throttle lever input. Thus, the turbine engine control system can control the turbine engine and the remote propulsor based on the single throttle lever input. In this way, the propulsion system of the present disclosure requires only a single throttle lever to be actuated by the pilot to control both the turbine engine and the remote propulsor, thus, reducing pilot workload, allowing the pilot to better focus attention on other needs. The remote propulsor can include at least one of an electrical propulsor (e.g., powered by an electric machine) or a second turbine engine (e.g., turboprop engine). The present disclosure provides for communication between the turbine engine and the remote propulsor via a data communication bus using at least one of wireless (e.g., 5G), the aircraft communication bus (e.g., ARINC), or a dedicated wire harness between the turbine engine and the remote propulsor.

In certain embodiments, the propulsion system disclosed herein is a hybrid electric propulsion system in which the turbine engine (e.g., turboprop engine) includes the electric machine and the propulsor can be electrically powered by the electric machine. The remote propulsor can also be powered by an electric machine. The electric machines in the propulsion system can operate at a high-speed requiring a reduction gearbox between the electric machine and the propulsors. For example, the propulsor of the turbine engine and the remote propulsor both include a reduction gearbox between the electric machine of the respective propulsor and the propulsor.

The reduction gearbox requires lubrication when the propulsor (e.g., the propulsor of the turbine engine and/or the remote propulsor) is operated, with the local turbine engine in a shutdown/standby condition, such that the lubrication system of the local turbine engine is not functional. In particular, in current turbine engines, a pump of the lubrication system is powered by the turbine shaft (e.g., the LP shaft or the HP shaft), and, thus, the lubrication system is unable to provide lubricant to the gearbox when the turbine engine is shut down (the turbine shaft stops rotating). Therefore, the present disclosure also provides for a hybrid electric turbine engine system that incorporates a reduction gearbox with a gearbox lubrication system that is separate from the engine lubrication system. The gearbox lubrication system comprises a pump and a lubricant reservoir (e.g., a sump). In some embodiments, the gearbox lubrication system includes a heat exchanger, a pressure sensor, or a level sensor. The pressure sensor and level sensor output signals can be communicated to the remote propulsor either by discrete outputs or by a digital data bus via the turbine engine control system. In one embodiment, the lubrication system supplies the lubricant to one or more engine bearings from the propulsor gearbox lubrication system, while the turbine engine is shut down, and the propulsor is electrically driven. The gearbox lubrication system is powered by a pump that is drivingly coupled to at least one of the propulsor shaft, the turbine shaft, or the gears of the gearbox. In this way, the lubrication system is a mechanically driven lubrication system.

The present disclosure also provides for an electrically driven lubrication system. The lubrication system allows the turbo-engine to be shut down, while still providing lubrication to the propulsor. The lubrication system can be powered by aircraft power or by electric power from the remote turbine engine, when the remote turbine engine is in control. The lubrication system can include lubricant sensors that provide data to the remote turbine engine as well as to the local turbine engine, so as to allow for safe control by the remote turbine engine.

The present disclose further provides a plurality of lubricant sensors (e.g., temperature sensors, pressure sensors, level sensors, or flow rate sensors) associated with the lubrication system that enable fault monitoring by the remote turbine engine. The turbine engine control system identifies failure modes of the lubricant sensors and operates the remote turbine engine at a reduced power or shutting down the propulsor based on the failure modes.

FIG.9is a schematic view of a propulsion system150having a turbine engine control system200, according to the present disclosure. The propulsion system150ofFIG.9is similar to the systems described above with respect toFIGS.2-5. As such, common elements of the description will not be repeated here.

The turbine engine10also includes a fuel system180for providing the fuel to the combustor26. For example, the fuel system180can include a fuel tank for storing the fuel, one or more fuel lines in flow communication with the fuel tank and the combustor26, and a fuel pump for delivering the fuel from the fuel tank to the combustor26through the one or more fuel lines. The turbine engine10includes a first propulsor38ahaving a plurality of first propulsor blades40aand a first propulsor shaft45a, and a first electric machine74a.

The propulsion system150also includes a second propulsor38bhaving a plurality of second propulsor blades40b. The second propulsor38bis remote (e.g., separate) from the turbine engine10. The second propulsor38bincludes a second propulsor shaft45bthat is drivingly coupled to a second electric machine74b. The second electric machine74bis in electrical communication with the first electric machine74avia a power conduit75. In this way, the first electric machine74aprovides power to the second electric machine74bvia the power conduit75. In some embodiments, the second electric machine74bprovides power to the first electric machine74avia the power conduit75. In some embodiments, the first electric machine74aand the second electric machine74bexchange power via a common energy storage device (e.g., one or more batteries). The second electric machine74bis operated as a motor and drives the plurality of second propulsor blades40b. The second electric machine74bcan be used to provide additional thrust output beyond that provided by the plurality of first propulsor blades40adriven by the turbine engine10.

The turbine engine control system200includes a controller202and a single input device, also referred to as a single throttle lever204. The controller202is in two-way communication with the propulsion system150(e.g., the turbine engine10, the second propulsor38b, and the single throttle lever204) for controlling aspects of the propulsion system150(e.g., the turbine engine10, the second propulsor38b, and the single throttle lever204). The controller202, or components thereof, may be located onboard the turbine engine10, onboard the aircraft, or can be located remote from each of the turbine engine10and the aircraft (e.g., on the ground or on a remote propulsor). The controller202can be a Full Authority Digital Engine Control (FADEC) that controls aspects of the turbine engine10.

The controller202may be a standalone controller or may be part of an engine controller to operate various systems of the turbine engine10. In this embodiment, the controller202is a computing device having one or more processors and a memory. The one or more processors can be any suitable processing device, including, but not limited to, a microprocessor, a microcontroller, an integrated circuit, a logic device, a programmable logic controller (PLC), an application specific integrated circuit (ASIC), or a Field Programmable Gate Array (FPGA). The memory can include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, a computer readable non-volatile medium (e.g., a flash memory), a RAM, a ROM, hard drives, flash drives, or other memory devices.

The memory can store information accessible by the one or more processors, including computer-readable instructions that can be executed by the one or more processors. The instructions can be any set of instructions or a sequence of instructions that, when executed by the one or more processors, cause the one or more processors and the controller202to perform operations. The controller202and, more specifically, the one or more processors are programmed or configured to perform these operations, such as the operations discussed further below. In some embodiments, the instructions can be executed by the one or more processors to cause the one or more processors to complete any of the operations and functions for which the controller202is configured, as will be described further below. The instructions can be software written in any suitable programming language or can be implemented in hardware. Additionally, or alternatively, the instructions can be executed in logically or virtually separate threads on the processors. The memory can further store data that can be accessed by the one or more processors.

The single throttle lever204is a single, pilot-controllable, power control lever for controlling a power level of the turbine engine10and the second propulsor38b. In this way, the turbine engine control system200includes a single throttle lever204(e.g., one throttle lever) for controlling aspects of both the turbine engine10and the second propulsor38b. The single throttle lever204may be a pilot-accessible mechanical lever, for instance, located in a cockpit or a flight deck of an aircraft, that allows the pilot to provide input to control the turbine engine control system200by physically moving the single throttle lever204through a predetermined physical range of motion corresponding to the operation range of the turbine engine control system200. The single throttle lever204physical range of motion may additionally include a first portion or a first range, wherein moving the single throttle lever204into the first portion corresponds to a forward movement of the aircraft, and a second portion or a second range, wherein moving the single throttle lever204into the second portion corresponds to a rearward movement of the aircraft.

The single throttle lever204can include a positional sensor that detects or converts the position of the single throttle lever204, relative to the operational range of the single throttle lever204motion (in both the first position and the second position), and generate an output signal (e.g., a control signal) indicative of the position of the single throttle lever204. For example, the single throttle lever204can include a lever anchored at a lever pivot, and rotatable over 88°, wherein the foremost 44° of rotation (relative to the pilot) may provide the first portion in the range of motion, while the rearmost 44° may provide the second portion, or reverse, range of motion. The positional sensor, in turn, generates a control signal that ranges from −44° to +44°. The 0° to 44° range defines the first position (corresponding to forward movement) and the 0° to −44° range defines the second position (corresponding to rearward movement). The single throttle lever204can be rotatable over any range, as desired, for generating a control signal for controlling the turbine engine10or portions thereof.

While a lever is described for the single throttle lever204, additional input devices allowing for pilot input can be used, such as a dial or a knob. Further, the single throttle lever204may be remote from the aircraft (e.g., remote controlled). Additionally, particular positions of the single throttle lever204may define different flight regimes, such as taxi, cruise, and takeoff. For example, a position of the single throttle lever204between 0 degrees and 20 degrees represents a taxi operation, a position of the single throttle lever204between 21 degrees and 30 degrees represents a cruise operation, and a position of the single throttle lever204at 44° represents a takeoff operation.

In operation, the turbine engine control system200receives an input (e.g., the control signal) from the single throttle lever204and controls the turbine engine10(e.g., at least one of the turbo-engine16or the first propulsor38a) and the second propulsor38bbased on the input from the single throttle lever204. In particular, the turbine engine control system200controls a rotational speed of the turbo-engine16and at least one of a rotational speed of the first propulsor38a, a pitch of the plurality of first propulsor blades40a, or a torque of the first propulsor38a. The turbine engine control system200also controls at least one of a rotational speed of the second propulsor38b, a pitch of the plurality of second propulsor blades40b, or a torque of the second propulsor38b. In some embodiments, the turbine engine control system200controls the fuel system80to increase or to decrease a fuel flow rate of the fuel to the combustor26to control the rotational speed of the core turbo-engine16. In some embodiments, the turbine engine control system200controls the first electric machine74ato control the first propulsor38a. For example, the controller202controls the first electric machine74aand the second electric machine74bto increase or to decrease an amount of power from the first electric machine74ato the first propulsor38afor controlling the rotational speed or the torque of the first propulsor38a. The turbine engine control system200similarly controls the second electric machine74bto control the second propulsor38b.

To control the turbine engine10, the first propulsor38a, and the second propulsor38b, the controller202receives a position input from the single throttle lever204. The controller202then converts the position input into at least one of a turbine engine setting or a first propulsor setting, and a second propulsor setting. The turbine engine setting includes a rotational speed of the turbo-engine16. The first propulsor setting includes at least one of a rotational speed of the first propulsor38a, a pitch of the plurality of first propulsor blades40a, or a torque of the first propulsor38a. The second propulsor setting includes at least one of a rotational speed of the second propulsor38b, a pitch of the plurality of second propulsor blades40b, or a torque of the second propulsor38b. The controller202then controls at least one of a rotational speed of the turbo-engine16, a rotational speed of the first propulsor38a, a pitch of the plurality of first propulsor blades40a, or a torque of the first propulsor38a, and at least one of a rotational speed of the second propulsor38b, a pitch of the plurality of second propulsor blades40b, or a torque of the second propulsor38bbased on the at least one of the turbine engine setting or the first propulsor setting, and the second propulsor setting. For example, the controller202controls the rotational speed of the turbo-engine16based on the turbine engine setting. The controller202controls at least one of the rotational speed of the first propulsor38a, the pitch of the plurality of first propulsor blades40a, or the torque of the first propulsor38abased on the first propulsor setting. The controller202controls at least one of the rotational speed of the second propulsor38b, the pitch of the plurality of second propulsor blades40b, or the torque of the second propulsor38bbased on the second propulsor setting.

In one embodiment, the controller202determines the turbine engine setting, the first propulsor setting, and the second propulsor setting based on one or more speed maps. For example, the controller202includes a turbo-engine speed map, a first propulsor speed map, and a second propulsor speed map. The turbine engine speed map, the first propulsor speed map, and the second propulsor speed map each operates as a demand map, or a look-up table, and may be configured with predetermined or dynamic profiles or setting values, for instance, via an implementation of an algorithm. In particular, the speed maps correlate various position inputs from the single throttle lever204to the turbine engine setting, the first propulsor setting, and the second propulsor setting. Each of the turbine engine speed map, the first propulsor speed map, and the second propulsor speed map may additionally define limitations to aircraft operation. For example, the turbine engine speed map, the first propulsor speed map, and the second propulsor speed map may provide limited maximum or minimum thrust or speed utilized during taxi, cruise, or takeoff. In another example, the turbine engine speed map, the first propulsor speed map, and the second propulsor speed map may define a maximum or a minimum turbine engine10speed, first propulsor38aspeed and pitch, or second propulsor38bspeed and pitch utilized during ground idle. In another example, the first propulsor speed map and the second propulsor speed map may define a constant propulsor speed with an adjusting propulsor blade pitch value. Alternatively, the first propulsor speed map and the second propulsor speed map may provide a constant propulsor blade pitch value with a varying propulsor speed. In yet another example, the first propulsor speed map and the second propulsor speed map may provide a varied speed or a varied pitch in either a forward thrust operation or a reverse thrust operation. In even yet another example, the first propulsor speed map and the second propulsor speed map may employ multiple maps, wherein, for instance, a forward control signal utilizes a forward propulsor speed map while a reverse control signal utilizes a reverse propulsor speed map. In even yet another example, the turbine engine speed map, the first propulsor speed map, and the second propulsor speed map may be configured to prevent excess or undesirable temperature conditions on the turbo-engine16and over-torque or overspeed on the first propulsor38aand the second propulsor38b, respectively.

The turbine engine speed map, the first propulsor speed map, and the second propulsor speed map may additionally define profiles of operation, wherein, for example, the turbine engine speed map, the first propulsor speed map, and the second propulsor speed map are configured to provide setting values for optimal aircraft efficiency, turbine engine efficiency, or propulsor efficiency. For example, the first propulsor speed map and the second propulsor speed map may be configured to provide an optimized propulsor speed for a corresponding turbine engine speed map value. The optimized profiles may additionally take into account additional sensor values, for instance, relating to air speed, altitude, etc., to provide an optimal turbine engine10speed, an optimal first propulsor38aspeed and first propulsor blade40apitch, and an optimal second propulsor38bspeed and second propulsor blade40bpitch, for example, for an optimized fuel efficiency profile, an optimized thrust profile, an optimized audible profile (i.e., by controlling pitch of the first propulsor blades40aor the second propulsor blades40b), or an optimized flight time efficiency profile. The optimization profiles may be further delineated by a given control signal from the single throttle lever204. The optimized profiles may be user-selectable by the pilot.

Additionally, or alternatively, at least one of the turbine engine speed map, the first propulsor speed map, or the second propulsor speed map may be configured to provide for a linear thrust response, that is, a linear thrust relative to the position of the single throttle lever204. The thrust response may be linear in a forward thrust operation or a reverse thrust operation. In this sense, “linear” may refer to the overall thrust which, as explained above, is affected by the configurations of the turbine engine10, the first propulsor38a, the first propulsor blades40a, the second propulsor38b, and the second propulsor blades40b. Thus, a linear thrust may not be proportionally related to, for example, only a change in one component (e.g., the core turbo-engine16, the first propulsor38a, the first propulsor blades40a, the second propulsor38b, or the second propulsor blades40b) of the configuration. Additionally, “linear” may define a maximum allowable change in setting values, regardless of how much change or how quickly a change was requested via the single throttle lever204. In this instance, a “linear” thrust response may only allow for a 1% increase in turbine engine power per second. In another sense, “linear” may define a feedback response from a perspective of the pilot. For instance, a pilot may define a feeling of “linear” thrust response as, for example, a 1% increase in thrust for one second, followed by a 2% increase in thrust for three seconds. Further, this thrust response may be different depending on the status, or flight regime, of the aircraft (e.g., taxi, cruise, takeoff, etc.).

Each of the turbine engine speed map, the first propulsor speed map, or the second propulsor speed map receive the control signal from the single throttle lever204, and the controller202determines the turbine engine setting (e.g., the turbo-engine setting and the first propulsor setting) and the second propulsor setting in response to, or as a function of, the received input. In some embodiments, the controller202determines a desired turbine engine setting (e.g., a desired turbo-engine setting and a desired first propulsor setting) and a desired second propulsor setting based on the received input from the single throttle lever204. The controller202then compares the desired turbine engine setting and the desired second propulsor setting to an actual turbine engine setting and an actual second propulsor setting. For example, the controller202determines the actual rotational speed of the turbo-engine16, the actual rotational speed of the first propulsor38aand the second propulsor38b, the actual pitch of the first propulsor blades40aand the second propulsor blades40b, and the actual torque of the first propulsor38aand the second propulsor38b.

The controller202then controls the turbo-engine16, the first propulsor38a, and the second propulsor38bbased on the comparison. For example, to speed up the first propulsor38aor the second propulsor38b, the controller202generates a positive value, that, in turn, decreases the first propulsor38apitch or the second propulsor38bpitch, thus increasing the rotational speed of the first propulsor38aor the second propulsor38b. Likewise, to slow down the first propulsor38aor the second propulsor38b, the controller202generates a negative value, that, in turn, increases the first propulsor38apitch or the second propulsor38bpitch, thus decreasing the rotational speed of the first propulsor38aor the second propulsor38b. In instances when the control signal is static, the operation of speeding up and/or slowing down of the first propulsor38aor the second propulsor38bwill ultimately equalize at a point wherein the controller202generates a neutral value, at which the pitch of the first propulsor38aor the second propulsor38bis neither increased nor decreased. The controller202can similarly control the turbo-engine16, the first propulsor38a, and the second propulsor38bduring a reverse thrust operation.

FIG.10is a schematic view of the propulsion system150having the turbine engine control system200, according to the present disclosure. As shown inFIG.10, the turbine engine control system200further includes an aircraft power supply206for powering the second electric machine74b. The aircraft power supply206includes at least one of a fuel cell, one or more batteries, or an auxiliary power unit (APU) of the aircraft. The aircraft power supply206can also power the first electric machine74a. In some embodiments, the second electric machine74breceives power from both the first electric machine74aand the aircraft power supply206. For example, the aircraft power supply206supplies the power during takeoff or landing, and the first electric machine74asupplies the power during cruise.

FIG.11is a schematic view of a propulsion system300having a turbine engine control system400, according to another embodiment. The turbine engine control system400is substantially similar to the turbine engine control system200ofFIGS.9and10. The same or similar reference numerals will be used for components of the turbine engine control system400that are the same as or similar to the components of the turbine engine control system200discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system300includes a plurality of turbine engines10including a first turbine engine10aand a second turbine engine10b. The first turbine engine10ahas a first turbo-engine16aand the second turbine engine10bhas a second turbo-engine16b. The second propulsor38bis drivingly coupled to the second turbo-engine16b(e.g., via the LP shaft36of the second turbo-engine16b) such that the second turbine engine10bincludes the second propulsor38b. In this way, the second turbo-engine16bpowers the second propulsor38b. The first turbine engine10aincludes a first fuel system180aand the second turbine engine10bincludes a second fuel system180b.

The turbine engine control system400includes one or more controllers402and one or more single throttle levers404including a first single throttle lever404aand a second single throttle lever404b. The turbine engine control system400also includes an aircraft power supply406. The one or more controllers402include a first controller402afor controlling the first turbine engine10aand a second controller402bfor controlling the second turbine engine10b. The first single throttle lever404ais in communication with the first controller402a. The second single throttle lever404bis in communication with the second controller402b. In this way, the first single throttle lever404acontrols the first turbine engine10aand the second single throttle lever404bcontrols the second turbine engine10b. The turbine engine control system400also includes a data communication bus405for providing communication between the first controller402aand the second controller402b. The data communication bus405can include wired communication (e.g., a wire harness, an aircraft data bus, such as ARINC, etc.) or wireless communication (e.g., Wi-Fi®, Bluetooth®, cellular communication, such as 5G, etc.). In this way, the first controller402acan control aspects of the second turbine engine10band the second controller402bcan control aspects of the first turbine engine10a, under certain operating conditions, as detailed further below.

In operation, the turbine engine control system400is substantially similar to the turbine engine control system200ofFIG.2. The first controller402acontrols the first turbine engine10a(e.g., the first turbo-engine16aand the first propulsor38a) based on the input from the first single throttle lever404a. The second controller402bcontrols the second turbine engine10b(e.g., the second turbo-engine16band the second propulsor38b) based on the input from the second single throttle lever404b.

The first controller402areceives operating mode signals indicative of an operating mode of the second turbine engine10bfrom the second controller402b. In some instances, the second turbine engine10benters a standby operating mode or a failure mode occurs. For example, the rotational speed of the second turbo-engine16breduces to less than a speed threshold for the current flight condition, or the first controller402aloses communication with the second controller402b. The standby operating mode can comprise at least one of a normal idling mode, an assisted idling mode, a banking mode, or a stopping mode. The normal idling mode occurs when the combustor26is operating (e.g., ignited) and the LP shaft36rotates at a rotational speed in a range of 45% to 70% of a nominal speed of the LP shaft36. The nominal speed is a speed of the LP shaft36at a maximum cruise condition of the second turbine engine10b(or the first turbine engine10a). The assisted idling mode occurs when the combustor26is operating and the LP shaft36rotates in a mechanically-assisted manner (e.g., via the starter motor82) at a rotational speed in a range of 20% to 60% of the nominal speed. The banking mode occurs when the combustor26is shut down (e.g., extinguished) and the LP shaft36rotates in a mechanically-assisted manner at a rotational speed in a range of 5% to 30% of the nominal speed. The stopping mode occurs when the combustor26is shut down and the LP shaft36stops rotating or windmills due to airflow through the turbine engine10.

The first controller402adetects that the second turbine engine10bis in the standby operating mode (e.g., based on the operating mode signals) and provides electrical power from the first turbine engine10a(e.g., the first electric machine74a) to the second turbine engine10b(e.g., the second electric machine74b) such that the first turbine engine10apowers the second turbine engine10b. During the standby operating mode of the second turbine engine10b, the first controller402acontrols the second turbine engine10band the second propulsor38bbased on the input from the first single throttle lever404a. In this way, the first controller402acontrols the rotational speed of the first turbo-engine16aand the second turbo-engine16b, the rotational speed and the torque of the first propulsor38aand the second propulsor38b, and the pitch of the first propulsor blades40aand the second propulsor blades40bbased on the input from the first single throttle lever404a. In some embodiments, the first controller402acontrols the second turbine engine10bbased on a predetermined position or a positional threshold of the first single throttle lever404a. For example, the pilot could move the first single throttle lever404ainto the predetermined position or beyond the positional threshold during any operating mode to control the second turbine engine10bwith the first controller402abased on the first single throttle lever404a.

FIG.12is a schematic view of a propulsion system500, according to another embodiment. The propulsion system500includes the turbine engine10. The turbine engine10includes a lubrication system600having a gearbox lubrication system602for the gearbox assembly46. The gearbox lubrication system602is entirely contained within the gearbox assembly46, as detailed further below.

FIG.13is an enlarged schematic view of the lubrication system600for the propulsion system500, according to the present disclosure. As shown inFIG.13, the turbine engine10includes one or more engine bearings37that support rotation of the HP shaft34, the LP shaft36, or other rotating components of the turbine engine10. The gearbox assembly46includes a gearbox housing47and a gear assembly650disposed within the gearbox housing47. The propulsor shaft45is drivingly coupled to the LP shaft36through the gear assembly650. The gear assembly650includes one or more gears652. InFIG.13, the gear assembly650is a planetary gear assembly that includes a first gear652a, one or more second gears652b, and a third gear652c. The first gear652ais a sun gear, the one or more second gears652bare planet gears, and the third gear52cis a ring gear. The LP shaft36is coupled to the first gear652aand the propulsor shaft45is coupled to the third gear652c. The gear assembly650can include any number of gears652in any configuration for transmitting power from the LP shaft36to the propulsor shaft45, as desired.

The gearbox lubrication system602includes a gearbox lubricant reservoir604, a pump606, and a gearbox lubricant supply line608. The gearbox lubricant reservoir604stores a lubricant therein. The lubricant can include any type of lubricant for lubricating the gearbox assembly46. For example, the lubricant is oil, or the like. The gearbox lubricant reservoir604includes at least one of a lubricant tank or a sump. The sump collects the lubricant that drains from the gear assembly650. The gearbox lubricant supply line608is in fluid communication with the gearbox lubricant reservoir604and the gear assembly650for supplying the lubricant from the gearbox lubricant reservoir604to the gear assembly650. The pump606is in fluid communication with the gearbox lubricant supply line608to pump the lubricant from the gearbox lubricant reservoir604to the gear assembly650through the gearbox lubricant supply line608. The pump606is drivingly coupled to the propulsor shaft45. In this way, the pump606pumps the lubricant from the gearbox lubricant reservoir604when the propulsor shaft45rotates. Thus, the lubrication system600is a mechanically-driven lubrication system. In some embodiments, the pump606is drivingly coupled to the LP shaft36. In some embodiments, the pump606is drivingly coupled to at least one of the plurality of gears652of the gear assembly650. Thus, the pump606is drivingly coupled to at least one of the propulsor shaft45, the LP shaft36, or the gear assembly650. The gearbox lubrication system602also includes a heat exchanger610in fluid communication with the gearbox lubricant supply line608. The heat exchanger610cools the lubricant as the lubricant flows through the gearbox lubricant supply line608before being supplied to the gear assembly650.

The gearbox lubrication system602also includes one or more lubricant sensors612for sensing information about the lubricant in the gearbox lubrication system602. The one or more lubricant sensors612are in communication with the controller202for sending the information about the lubricant to the controller202. For the propulsion system300(FIG.11), the one or more lubricant sensors612send the information about the lubricant to the first controller402a(FIG.11) of the first turbine engine10a(FIG.11) and to the second controller402b(FIG.11) of the second turbine engine10b(FIG.11). In this way, both the local engine (e.g., the first turbine engine10a) and the remote engine (e.g., the second turbine engine10b) receive the information about the lubricant from the one or more lubricant sensors612of each of the turbine engines10a,10b.

The one or more lubricant sensors612include a lubricant level sensor612aand a lubricant pressure sensor612b. The lubricant level sensor612asenses a level of the lubricant in the gearbox lubricant reservoir604. The lubricant pressure sensor612bsenses a pressure of the lubricant in the gearbox lubricant supply line608. The lubricant sensors612can also include a lubricant temperature sensor for sensing a temperature of the lubricant, and a lubricant flow sensor for sensing a flow rate of the lubricant. The propulsion system500can utilize the information from the sensors to detect a failure mode of the lubrication system600and to control the turbine engine10in response to the failure mode, as detailed further below.

The lubrication system600also includes an engine lubrication system620that supplies lubricant to the one or more engine bearings37. The gearbox lubrication system602is fluidly separate from the engine lubrication system620. In this way, the gearbox lubrication system602lubricates the gear assembly650without receiving lubricant from the engine lubrication system620. The engine lubrication system620includes an engine lubricant supply line622and an engine lubricant return line624. Although not shown, the engine lubrication system620also includes a pump and a gearbox lubricant reservoir. The engine lubricant supply line622is in fluid communication with the engine bearings37for supplying the lubricant to the engine bearings37through the engine lubricant supply line622. The engine lubricant return line624is in fluid communication with the engine bearings37for returning the lubricant from the engine bearings37back to the engine lubrication system620.

The lubrication system600includes a gearbox engine bearing supply line630and a gearbox engine bearing return line632. The gearbox engine bearing supply line630is in fluid communication with the gearbox lubricant supply line608and the engine bearings37for supplying the lubricant from the gearbox lubrication system602to the engine bearings37. The gearbox engine bearing return line632is in fluid communication with the engine bearings37and the gearbox lubricant reservoir604for returning the lubricant from the engine bearings37to the gearbox lubricant reservoir604.

The lubrication system further includes one or more valves640in fluid communication with the engine lubricant supply line622, the gearbox engine bearing supply line630, the engine lubricant return line624, and the gearbox engine bearing return line632. The one or more valves640can be shuttle valves for allowing the lubricant to flow, or preventing the lubricant from flowing, through the engine lubricant supply line622, the gearbox engine bearing supply line630, the engine lubricant return line624, and the gearbox engine bearing return line632. The one or more valves640can include any type of valve, such as, for example, check valves, for allowing the lubricant to flow, or preventing the lubricant from flowing, through the engine lubricant supply line622, the gearbox engine bearing supply line630, the engine lubricant return line624, and the gearbox engine bearing return line632. The one or more valves640include a first valve640ain fluid communication with the engine lubricant supply line622and the gearbox engine bearing supply line630. The first valve640aallows the lubricant to flow, or prevents the lubricant from flowing, through the engine lubricant supply line622and the gearbox engine bearing supply line630. The one or more valves640include a second valve640bin fluid communication with the engine lubricant return line624and the gearbox engine bearing return line632. The second valve640ballows the lubricant to flow, or prevents the lubricant from flowing, through the engine lubricant return line624and the gearbox engine bearing return line632.

The engine lubrication system620communicates an engine lubricant pressure signal621to the one or more valves640for actuating the one or more valves640, as detailed further below. In some embodiments, the one or more valves640are electrically actuated by the turbine engine control system400. For example, the controller202controls the one or more valves640to actuate the one or more valves640.

In operation, the gearbox lubrication system602supplies the lubricant to the gear assembly650to lubricate the gear assembly650. The gearbox lubrication system602is a standalone lubrication system that operates to supply the lubricant to the gear assembly650entirely separate from the engine lubrication system620. In particular, the pump606pumps the lubricant from the gearbox lubricant reservoir604to the gear assembly650through the gearbox lubricant supply line608. The heat exchanger610cools the lubricant in the gearbox lubricant supply line608as the lubricant flows through the gearbox lubricant supply line608and to the gear assembly650. After lubricating the gear assembly650(e.g., the plurality of gears652), the lubricant drains into the gearbox lubricant reservoir604(e.g., the sump). The lubricant is then re-circulated through the gearbox lubrication system602. Thus, the gearbox lubrication system602is a self-contained lubrication system for supplying the lubricant to the gear assembly650without the use of the engine lubrication system620. In this way, the gearbox lubrication system602can operate even if the turbo-engine16(FIG.12) shuts down.

During normal operation, the engine lubrication system620supplies the lubricant to the engine bearings37through the engine lubricant supply line622. During the normal operation, the pressure of the lubricant in the engine lubrication system620(e.g., the engine lubricant supply line622) is greater than the pressure of the lubricant in the gearbox lubrication system602(e.g., the gearbox engine bearing supply line630). Thus, the first valve640aopens the engine lubricant supply line622and closes the gearbox engine bearing supply line630. In this way, the lubricant flows through the engine lubricant supply line622to the engine bearings37for lubricating the engine bearings37, and the first valve640aprevents the lubricant in the gearbox lubrication system602from flowing to the engine bearings37. After lubricating the engine bearings37, the engine lubricant return line624directs the lubricant back through the engine lubrication system620. In particular, the second valve640bopens the engine lubricant return line624and closes the gearbox engine bearing return line632. In this way, the lubricant flows from the engine bearings37to the engine lubrication system620through the engine lubricant return line624to recirculate the lubricant to the engine bearings37.

The gearbox lubrication system602supplies the lubricant to the engine bearings37when the turbine engine10is shut down. The engine lubrication system620reduces the supply, or stops the supply, of the lubricant to the engine bearings37when the turbine engine10is shut down such that the pressure of the lubricant in the engine lubrication system620reduces. Accordingly, the pressure of the lubricant in the gearbox lubrication system602is greater than the pressure of lubricant in the engine lubrication system620when the turbine engine10is shut down. The first valve640acloses the engine lubricant supply line622and opens the gearbox engine bearing supply line630. In this way, the lubricant from the gearbox lubrication system602flows through the gearbox engine bearing supply line630to the engine bearings37to lubricate the engine bearings37. After lubricating the engine bearings37, the gearbox engine bearing return line632directs the lubricant back through the gearbox lubrication system602. In particular, the second valve640bopens the gearbox engine bearing return line632and closes the engine lubricant return line624. In this way, the lubricant flows from the engine bearings37to the gearbox lubricant reservoir604to recirculate the lubricant through the gearbox lubrication system602.

FIG.14is a schematic view of a propulsion system700having a lubrication system800, according to another embodiment. The lubrication system800is substantially similar to the lubrication system600ofFIGS.12and13. The same or similar reference numerals will be used for components of the lubrication system800that are the same as or similar to the components of the lubrication system600discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system700includes a first propulsor738a, a second propulsor738b, and an electric power supply715. The first propulsor738aincludes a plurality of first propulsor blades740aand a first propulsor shaft745athat is drivingly coupled to a first electric machine770a. In particular, the first electric machine770aincludes a first electric machine shaft771a, and the first propulsor shaft745ais drivingly coupled to the first electric machine shaft771a. The first propulsor738aalso includes a first gearbox assembly746a, and the first propulsor shaft745ais drivingly coupled to the first electric machine shaft771athrough the first gearbox assembly746a. The first gearbox assembly746aincludes a first gearbox lubrication system802athat is a self-contained lubrication system similar to the gearbox lubrication system602ofFIGS.12and13. In some embodiments, the first electric machine770aincludes an electric machine lubrication system, such as the electric machine lubrication system1200, detailed below with respect toFIGS.16and17.

The second propulsor738bincludes a plurality of second propulsor blades740band a second propulsor shaft745bthat is drivingly coupled to a second electric machine770b. In particular, the second electric machine770bincludes a second electric machine shaft771b, and the second propulsor shaft745bis drivingly coupled to the second electric machine shaft771b. The second propulsor738balso includes a second gearbox assembly746b, and the second propulsor shaft745bis drivingly coupled to the second electric machine shaft771bthrough the second gearbox assembly746b. The second gearbox assembly746bincludes a second gearbox lubrication system802bthat is a self-contained lubrication system similar to the gearbox lubrication system602ofFIGS.12and13. In some embodiments, the second electric machine770bincludes an electric machine lubrication system, such as the electric machine lubrication system1200, detailed below with respect toFIGS.16and17.

The electric power supply715includes a turbo-engine716. The turbo-engine716includes a third electric machine770cthat is drivingly coupled to the LP shaft36. In this way, the turbo-engine716supplies mechanical power to the third electric machine770c, and the third electric machine770cconverts the mechanical power into electric power. The electric power supply715is in electrical communication with the first electric machine770avia a first electric conduit775a, and with the second electric machine770bvia a second electric conduit775b. In this way, the electric power supply715supplies electric power to the first electric machine770aand to the second electric machine770bfor powering the first propulsor738aand the second propulsor738b, respectively. In particular, the first electric machine770aand the second electric machine770bare in electrical communication with the third electric machine770c, and the third electric machine770csupplies the electric power to the first electric machine770aand to the second electric machine770b. In some embodiments, the electric power supply715includes a fuel cell, a battery, or an APU of the aircraft. In this way, the electric power supply715includes at least one of the turbo-engine716, a fuel cell, a battery, or an APU.

FIG.15is a schematic view of a propulsion system900, according to another embodiment. The propulsion system900includes a turbine engine910including a lubrication system1000having a gearbox lubrication system1002. The turbine engine910includes an electric machine970drivingly coupled to the LP shaft36. The electric machine970is positioned forward of the turbine section27, and, particularly, is positioned forward of the LP turbine30. In this way, the electric machine970is positioned axially between the gearbox assembly46and the turbine section27. The turbine engine910also includes a clutch973that engages the electric machine970to the LP shaft36and disengages the electric machine970from the LP shaft36. In particular, the clutch973engages the electric machine970to the LP shaft36when the electric machine970operates in an electric generator mode to generate electric power. The clutch973disengages the electric machine970from the LP shaft36when the electric machine970operates in an electric motor mode to supply the electric power to the propulsor38(e.g., through the gearbox assembly46).

FIG.16is a schematic view of a propulsion system1100, according to another embodiment. The propulsion system1100includes the turbine engine10, the first propulsor38a, and the second propulsor38b. The propulsion system1100includes an electric machine lubrication system1200for the second electric machine74b. The electric machine lubrication system1200is entirely contained within the second electric machine74b, as detailed further below. In some embodiments, the propulsion system1100includes a fluid conduit1175that fluidly couples the second electric machine74bwith the first electric machine74a. In such embodiments, the electric machine lubrication system1200supplies the lubricant from the second electric machine74bto the first electric machine74avia the fluid conduit1175. In some embodiments, the second electric machine74breceives the lubricant from the first electric machine74a. In some embodiments, the first electric machine74aincludes the electric machine lubrication system1200.

FIG.17is an enlarged schematic internal view of the second electric machine74bhaving the electric machine lubrication system1200, according to the present disclosure. As shown inFIG.17, the second electric machine74bincludes one or more electric machine bearings1290that support rotation of the second propulsor shaft45b. The one or more electric machine bearings1290include one or more first electric machine bearings1290aand one or more second electric machine bearings1290b. The second electric machine74bincludes a rotor1292and a stator1294. The rotor1292is coupled to the second propulsor shaft45band rotates with rotation of the second propulsor shaft45b. In this way, the rotor1292rotates with respect to the stator1294, generating electrical power. The stator1294includes one or more electrical coils1295. When the second electric machine74boperates as an electric generator, the second electric machine74bgenerates electric power due to a movement of the rotor1292relative to the stator1294(e.g., the electric coils1295) in a first direction. When the second electric machine74boperates as a motor, the second electric machine74bapplies a torque on the second propulsor shaft45bdue to the movement of the rotor1292relative to the stator1294in a second direction that is opposite the first direction.

The electric machine lubrication system1200includes a lubricant reservoir1202, an electric machine pump1204, a lubricant supply line1206, a lubricant return line1208, one or more lubricant injectors1210, and a heat exchanger1212. The lubricant reservoir1202stores the lubricant therein. In some embodiments, the lubricant reservoir1202is a sump within the second electric machine74b. The electric machine pump1204is drivingly coupled to the second propulsor shaft45b. The lubricant supply line1206is in fluid communication with the electric machine pump1204and the one or more lubricant injectors1210. The heat exchanger1212is in fluid communication with the lubricant supply line1206for cooling the lubricant in the lubricant supply line1206. The lubricant return line1208is in fluid communication with the one or more electric machine bearings1290and the lubricant reservoir1202for returning the lubricant from the electric machine bearings1290to the lubricant reservoir1202.

In operation, the electric machine pump1204pumps the lubricant from the lubricant reservoir1202to the one or more electric machine bearings1290through the lubricant supply line1206. In particular, the one or more lubricant injectors1210inject the lubricant from the lubricant supply line1206to the one or more electric machine bearings1290. The heat exchanger1212cools the lubricant in the lubricant supply line1206as the lubricant flows through the heat exchanger1212. In particular, the volume of air50(e.g., cold air) flows through the heat exchanger and heat from the lubricant in the heat exchanger1212is transferred to the volume of air50as the lubricant flows through the heat exchanger1212. The lubricant then drains through the lubricant return line1208into the lubricant reservoir1202such that the electric machine lubrication system1200recirculates the lubricant through the electric machine lubrication system1200.

FIG.18is a schematic view of a propulsion system1300, according to another embodiment. The propulsion system1300is substantially similar to the propulsion system1200ofFIG.12. The same or similar reference numerals will be used for components of the propulsion system1300that are the same as or similar to the components of the propulsion system500discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system1300includes a turbine engine1310. The turbine engine1310is a turboprop engine in a pusher configuration. The turbine engine1310includes a propulsor section1314, a turbo-engine1316, and an exhaust section1332. The turbo-engine1316includes a compressor1322, a combustor1326, and a turbine section1327having an HP turbine1328and an LP turbine1330. The turbine engine1310also includes an HP shaft1334and an LP shaft1336. The propulsor section1314is arranged aft of the turbo-engine1316in the pusher configuration. The propulsor section1314includes a propulsor1338having a plurality of propulsor blades1340coupled to a propulsor shaft1345. The propulsor shaft1345is drivingly coupled to the LP shaft1336through a gearbox assembly1346. The turbine engine1310includes a lubrication system1400. The gearbox assembly1346includes a gearbox lubrication system1402.

The turbine engine1310also includes an electric machine1370that is drivingly coupled to the LP shaft1336through a clutch1371. The turbine engine1310further includes an accessory gearbox1372and a starter motor1374. The turbine engine1310operates substantially similar to the turbine engine10. The combustion gases68flow through the exhaust section1332under the propulsor1338, while bypass gas69is directed outside the engine.

FIG.19is a schematic view of a propulsion system1500, according to another embodiment. The propulsion system1500includes a turbine engine1510and a lubrication system1600for the turbine engine1510. The lubrication system1600includes a lubricant reservoir1602, a pump1604powered by a motor1606, a lubricant supply line1608, and a lubricant return line1610. The lubricant reservoir1602includes at least one of a sump or a tank, and stores the lubricant therein. The lubricant supply line1608is in fluid communication with the lubricant reservoir1602, the gearbox assembly46, and the one or more engine bearings37for supplying the lubricant from the lubricant reservoir1602to at least one of the gearbox assembly46or the one or more engine bearings37. The pump1604is in fluid communication with the lubricant reservoir1602and the lubricant supply line1608to pump. The lubrication system1600also includes a heat exchanger1618in fluid communication with the lubricant supply line1608. The heat exchanger1618cools the lubricant as the lubricant flows through the lubricant supply line1608before being supplied to the gearbox assembly46or the turbine engine10.

The lubrication system1600includes a gearbox lubrication system1612and an engine lubrication system1620. The gearbox lubrication system1612includes a gearbox lubricant supply line1614and a gearbox lubricant return line1616. The engine lubrication system1620includes one or more engine lubricant supply lines1622and one or more engine lubricant return lines1624. The one or more engine lubricant supply lines1622each includes a valve1626that open to allow the lubricant to flow through the engine lubricant supply lines1622and close to prevent the lubricant from flowing through the engine lubricant supply lines1622.

The pump1604is in fluid communication with the lubricant supply line1608to pump the lubricant from the lubricant reservoir1602to the at least one of the gearbox assembly46or the engine bearings37through the lubricant supply line1608. The pump1604is drivingly coupled to the motor1606. The motor1606is an electric motor that converts electric power into mechanical power through a motor shaft for powering the pump1604. In some embodiments, the motor1606is drivingly coupled to accessory gearbox assembly80(e.g., through a motor gear assembly1607) for generating electric power. In some embodiments, the motor1606receives electric power from an electric power supply1630. The electric power supply1630includes at least one of the aircraft (e.g., batteries, APU, fuel cell, etc.) or a second turbine engine (e.g., the second turbine engine10bofFIG.11). Thus, the motor1606receives electric power from at least one of the turbine engine1510(e.g., through the accessory gearbox assembly80), the aircraft, or the second turbine engine10b. In this way, the lubrication system1600is an electrically-driven lubrication system. The lubrication system1600also includes one or more lubricant sensors1640including a lubricant level sensor1640aand a lubricant pressure sensor1640b. The lubricant sensors1640can also include a lubricant temperature sensor for sensing a temperature of the lubricant, and a lubricant flow sensor for sensing a flow rate of the lubricant.

In operation, the lubrication system1600supplies the lubricant to the gearbox assembly46to lubricate the gear assembly650. In particular, the pump1604pumps the lubricant from the lubricant reservoir1602to the gearbox assembly46through the lubricant supply line1608and through the gearbox lubricant supply line1614. The heat exchanger1618cools the lubricant in the lubricant supply line1608as the lubricant flows through the lubricant supply line1608and to the gearbox assembly46. After lubricating the gear assembly650, the lubricant drains into the lubricant reservoir1602(e.g., the sump) through the lubricant return line1610and the gearbox lubricant return line1616. The lubricant is then re-circulated through the lubrication system1600. As discussed above, the pump1604is electrically driven by the motor1606. In this way, the lubrication system1600can operate to lubricate the gear assembly650even if the turbo-engine16shuts down or is operating in the standby operating mode and the propulsor38is electrically driven (e.g., by the electric machine74or by the second electric machine74bofFIG.11).

The lubrication system1600can also supply the lubricant to the engine bearings37through the engine lubricant supply lines1622. In particular, the lubrication system1600supplies the lubricant to the engine bearings37when the turbine engine1510is shut down or in the standby operating mode. The valve1626of each engine lubricant supply line1622opens to allow the lubricant to flow to the engine bearings37through the engine lubricant supply lines1622. After lubricating the engine bearings37, the engine lubricant return line1624directs the lubricant to the lubricant reservoir1602through the lubricant return line1610. In this way, the lubrication system1600recirculates the lubricant through the lubrication system1600. The valve1626of each engine lubricant supply line1622opens and closes based on a pressure of the lubricant in the engine lubrication system1620. In some embodiments, a controller (the controller202ofFIG.9) controls the valve1626to open and to close the valve1626.

FIG.20is a schematic view of a propulsion system1700, according to another embodiment. The propulsion system1700is substantially similar to the propulsion system1500ofFIG.19. The same or similar reference numerals will be used for components of the propulsion system1700that are the same as or similar to the components of the propulsion system1500discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system1700includes a turbine engine1710that is substantially similar to the turbine engine10(FIG.1). The turbine engine1710includes an electric machine1770that is drivingly coupled to the LP shaft36via a clutch1771. The electric machine1770is positioned forward of the turbo-engine16, and is between the gearbox assembly46and the turbine section27(e.g., the LP turbine30). The propulsion system1700also includes a lubrication system1800that is substantially similar to the lubrication system1600ofFIG.19.

FIG.21is a schematic view of a propulsion system1900, according to another embodiment. The propulsion system1900is substantially similar to the propulsion system1500ofFIG.19. The same or similar reference numerals will be used for components of the propulsion system1900that are the same as or similar to the components of the propulsion system1500discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system1900includes a turbine engine1910that is substantially similar to the turbine engine10(FIG.1). The propulsion system1900also includes a lubrication system2000that is substantially similar to the lubrication system1600ofFIG.19. The lubrication system2000includes a pump2004that is coupled to the accessory gearbox assembly80through a clutch2005. The pump2004is powered by a motor2006. The pump2004and the motor2006are disposed within the outer casing18(FIG.1) of the turbine engine1910. The clutch2005engages the pump2004to the accessory gearbox assembly80when the turbo-engine16is operating to reduce the electrical demand from the motor2006in powering the pump2004. In this way, the turbo-engine16powers the pump2004while the turbo-engine16is operating. The clutch2005disengages the pump2004from the accessory gearbox assembly80when the turbo-engine16is shut down. In this way, the motor2006powers the pump2004when the turbo-engine16is shut down, and the propulsor38is powered by the electric machine74.

FIG.22is a schematic view of a propulsion system2100, according to another embodiment. The propulsion system2100is substantially similar to the propulsion system300ofFIG.11. The same or similar reference numerals will be used for components of the propulsion system2100that are the same as or similar to the components of the propulsion system300discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system2100includes a plurality of turbine engines2110including a first turbine engine2110aand a second turbine engine2110b. The first turbine engine2110aincludes a first gearbox assembly46aincluding first gearbox lubrication system602ahaving one or more lubricant sensors612. The second turbine engine2110bincludes a second gearbox assembly46bincluding second gearbox lubrication system602bhaving one or more lubricant sensors612.

The propulsion system2100includes a turbine engine control system2200that is substantially similar to the turbine engine control system400ofFIG.11. The turbine engine control system2200includes one or more controllers2202and includes a first controller2202afor the first turbine engine2110aand a second controller2202bfor the second turbine engine2110b. The first controller2202ais in communication with the second controller2202bvia a data communication bus2205.

The first controller2202ais in communication with the lubricant sensors612of the first gearbox lubrication system602a. In particular, the first controller2202ais a dual channel controller and the lubricant sensors612of the first gearbox lubrication system602ainclude two communication outputs. Each communication output of the lubricant sensors612of the first gearbox lubrication system602ais in communication with a respective channel of the first controller2202a. In this way, the lubricant sensors612are dual redundant sensors (e.g., sensors having two channels to the first controller2202a).

Similarly, the second controller2202bis in communication with the lubricant sensors612of the second gearbox lubrication system602b. In particular, the second controller2202bis a dual channel controller and the lubricant sensors612of the second gearbox lubrication system602binclude two communication outputs. Each communication output of the lubricant sensors612of the second gearbox lubrication system602bis in communication with a respective channel of the second controller2202b. In this way, the lubricant sensors612are dual redundant sensors (e.g., sensors having two channels to the second controller2202b). Each lubricant sensor612includes discrete outputs that are electrically isolated from each other such that no single failure can occur in a loss of thrust control.

FIG.23is a schematic view of a propulsion system2300, according to another embodiment. The propulsion system2300is substantially similar to the propulsion system2100ofFIG.22. The same or similar reference numerals will be used for components of the propulsion system2300that are the same as or similar to the components of the propulsion system2100discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system2300includes a plurality of turbine engines2310including a first turbine engine2310aand a second turbine engine2310b. The first turbine engine2310aincludes a first gearbox assembly46aincluding first gearbox lubrication system602ahaving one or more lubricant sensors612. The second turbine engine2310bincludes a second gearbox assembly46bincluding second gearbox lubrication system602bhaving one or more lubricant sensors612.

The propulsion system2300includes a turbine engine control system2400that is substantially similar to the turbine engine control system2200ofFIG.22. The turbine engine control system2400includes one or more controllers2402and includes a first controller2402afor the first turbine engine2310aand a second controller2402bfor the second turbine engine2310b. The first controller2402ais in communication with the second controller2402bvia a data communication bus2405.

The first controller2402aand the second controller2402bare each dual channel controllers. The lubricant sensors612are quad redundant sensors (e.g., sensors having four channels) with individual outputs to the first controller2402aand the second controller2402b. In particular, the lubricant sensors612of the first gearbox lubrication system602aare in communication with the first controller2402aand the second controller2402b. The lubricant sensors612of the second gearbox lubrication system602bare in communication with the second controller2402band the first controller2402a. Each lubricant sensor612includes discrete outputs that are electrically isolated from each other such that no single failure can occur in a loss of thrust control. In this way, the lubricant sensors612of the first gearbox lubrication system602ainclude two channels to the first controller2402aand two channels to the second controller2402b. Similarly, the lubricant sensors612of the second gearbox lubrication system602binclude two channels to the second controller2402band two channels to the first controller2402a. Accordingly, each lubricant sensor612is a quad redundant sensor having four channels.

FIG.24is a schematic view of a propulsion system2500, according to another embodiment. The propulsion system2500is substantially similar to the propulsion system1500ofFIG.19. The same or similar reference numerals will be used for components of the propulsion system2500that are the same as or similar to the components of the propulsion system1500discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system2500includes a plurality of turbine engines2510including a first turbine engine2510aand a second turbine engine2510b. The first turbine engine2510aincludes a first lubrication system1600a, and the second turbine engine2510bincludes a second lubrication system1600b. The propulsion system2500also includes a turbine engine control system2600that is substantially similar to the turbine engine control system2200ofFIG.22. The turbine engine control system2600includes one or more controllers2602includes a first controller2602afor the first turbine engine2510aand a second controller2602bfor the second turbine engine2510b. The first controller2602ais in communication with the second controller2602bvia a data communication bus2605.

The first controller2602aand the second controller2602bare each dual channel controllers. The lubricant sensors1640are dual redundant sensors (e.g., sensors having two channels) with individual outputs. In particular, the lubricant sensors1640of the first lubrication system1600aare in communication with the first controller2602a. The lubricant sensors1640of the second lubrication system1600bare in communication with the second controller2602b. Each lubricant sensor1640includes discrete outputs that are electrically isolated from each other such that no single failure can occur in a loss of thrust control. In this way, the lubricant sensors1640of the first lubrication system1600ainclude two channels to the first controller2602a. Similarly, the lubricant sensors1640of the second lubrication system1600binclude two channels to the second controller2602b. Accordingly, each lubricant sensor1640is a dual redundant sensor having two channels.

FIG.25is a schematic view of a propulsion system2700, according to another embodiment. The propulsion system2700is substantially similar to the propulsion system2500ofFIG.24. The same or similar reference numerals will be used for components of the propulsion system2700that are the same as or similar to the components of the propulsion system2500discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here

The propulsion system2700includes a plurality of turbine engines2710including a first turbine engine2710aand a second turbine engine2710b. The first turbine engine2710aincludes a first lubrication system1600ahaving one or more lubricant sensors1640. The second turbine engine2710bincludes a second lubrication system1600bhaving one or more lubricant sensors1640.

The propulsion system2700includes a turbine engine control system2800that is substantially similar to the turbine engine control system2400ofFIG.23. The turbine engine control system2800includes one or more controllers2802and includes a first controller2802afor the first turbine engine2710aand a second controller2802bfor the second turbine engine2710b. The first controller2802ais in communication with the second controller2802bvia a data communication bus2805.

The first controller2802aand the second controller2802bare each dual channel controllers. The lubricant sensors1640are quad redundant sensors (e.g., sensors having four channels) with individual outputs to the first controller2802aand the second controller2802b. In particular, the lubricant sensors1640of the first lubrication system1600aare in communication with the first controller2802aand the second controller2802b. The lubricant sensors1640of the second lubrication system1600bare in communication with the second controller2802band the first controller2802a. Each lubricant sensor1640includes discrete outputs that are electrically isolated from each other such that no single failure can occur in a loss of thrust control. In this way, the lubricant sensors1640of the first lubrication system1600ainclude two channels to the first controller2802aand two channels to the second controller2802b. Similarly, the lubricant sensors1640of the second lubrication system1600binclude two channels to the second controller2802band two channels to the first controller2802a. Accordingly, each lubricant sensor1640is a quad redundant sensor having four channels.

FIG.26Ais a schematic view of a propulsion system2900with a lubrication system3000in a normal operation mode, according to another embodiment.FIG.26Bis a schematic view of the propulsion system2900with the lubrication system3000in a failure operation mode, according to another embodiment. The propulsion system2900is substantially similar to the propulsion system1500ofFIG.19. The same or similar reference numerals will be used for components of the propulsion system2900that are the same as or similar to the components of the propulsion system300discussed above, unless stated otherwise. The description of these components above also applies to this embodiment, and a detailed description of these components is omitted here.

The propulsion system2900includes a plurality of turbine engines2910including a first turbine engine2910aand a second turbine engine2910b. The lubrication system3000includes a first lubrication system3000a, a second lubrication system3000b, and an auxiliary lubrication system3050. The first lubrication system3000aand the second lubrication system3000bare substantially similar to the lubrication system1600.

The auxiliary lubrication system3050includes an auxiliary pump3052, an auxiliary motor3054that is drivingly coupled to the auxiliary pump3052by a pump shaft3056, an auxiliary lubricant supply line3058, and an auxiliary lubricant return line3060. The auxiliary pump3052is in fluid communication with auxiliary lubricant supply line3058and the auxiliary lubricant return line3060. The auxiliary lubricant supply line3058is in fluid communication with auxiliary pump3052and the first gearbox assembly46aand the second gearbox assembly46bfor supplying the lubricant to the gear assembly650of the first gearbox assembly46aand the second gearbox assembly46b. The auxiliary lubricant return line3060is in fluid communication with the first gearbox assembly46aand the second gearbox assembly46b, and the auxiliary pump3052for returning the lubricant from the first gearbox assembly46aand the second gearbox assembly46bto the auxiliary pump3052.

The auxiliary lubrication system3050includes an auxiliary lubricant valve3070in fluid communication with the auxiliary pump3052and the auxiliary lubricant supply line3058. The auxiliary lubricant valve3070is a shuttle valve, but can include any type of valve for allowing the lubricant to flow, or preventing the lubricant from flowing, through the auxiliary lubricant supply line3058based on a pressure of the lubricant in the engine lubrication system1620. The auxiliary lubricant valve3070includes a lubricant pressure flowpath3072, an auxiliary lubricant flowpath3074, and a valve element3076. The lubricant pressure flowpath3072is in fluid communication with a lubricant pressure signal line3077of the engine lubrication system1620for sending a lubricant pressure signal to the auxiliary lubricant valve3070. The auxiliary lubricant flowpath3074is in fluid communication with the auxiliary lubricant supply line3058. The valve element3076is disposed within the lubricant pressure flowpath3072and the auxiliary lubricant flowpath3074, and moves to allow or to block the lubricant in the lubricant pressure flowpath3072and the auxiliary lubricant flowpath3074, as detailed further below.

The lubricant pressure signal line3077includes a lubricant pressure signal line check valve3079disposed therein. The auxiliary lubricant supply line3058includes a supply line check valve3080disposed therein. The auxiliary lubricant return line3060includes one or more return line check valves3082disposed therein. In particular, the auxiliary lubricant return line3060includes a return line check valve3082in fluid communication with the first lubrication system3000a. The auxiliary lubricant return line3060includes a return line check valve3082in fluid communication with the second lubrication system3000b. The lubrication system3000also includes a pump check valve3084in fluid communication with the auxiliary lubricant return line3060and the auxiliary pump3052.

With reference toFIG.26A, the lubrication system3000operates the same as the lubrication system1600ofFIG.19. During normal operation, the lubrication system3000supplies the lubricant to the gear assembly650or the engine bearings37of the first turbine engine2910aand the second turbine engine2910b. The lubricant pressure signal line3077communicates the pressure of the lubricant in the lubrication system3000to the auxiliary lubricant valve3070. During the normal operation, the pressure of the lubricant in the lubrication system3000opens the lubricant pressure signal line check valve3079and forces the valve element3076to close the auxiliary lubricant flowpath3074to prevent the lubricant from flowing through the auxiliary lubrication system3050(e.g., through the auxiliary lubricant supply line3058).

With reference toFIG.26B, the auxiliary lubrication system3050supplies the lubricant to the lubrication system3000when the lubrication system3000fails or one of the plurality of turbine engines2910shuts down. In such instances, the pressure of the lubricant in the lubrication system3000reduces below a threshold such that the valve element3076overcomes the pressure in the lubricant pressure signal line3077. In this way, the valve element3076moves to open the auxiliary lubricant flowpath3074. The auxiliary motor3054powers the auxiliary pump3052to pump the lubricant through the auxiliary lubricant supply line3058. Thus, the pressure of the lubricant in the auxiliary lubricant supply line3058opens the supply line check valve3080and the auxiliary lubrication system3050supplies the lubricant to at least one of the gear assembly650or the engine bearings37through the auxiliary lubricant supply line3058. As the auxiliary pump3052continues to pump the lubricant, the lubricant drains from the at least one of the gear assembly650or the engine bearings37and flows through the auxiliary lubricant return line3060. Thus, the pressure of the lubricant in the auxiliary lubricant return line3060opens the one or more return line check valves3082open to allow the lubricant to flow back to the inlet of the auxiliary pump3052. The pressure of the lubricant from the auxiliary pump3052closes the pump check valve3084such that the lubricant flows through the auxiliary lubricant valve3070into the auxiliary lubricant supply line3058and the pump check valve3084prevents the lubricant from flowing from the auxiliary pump3052into the auxiliary lubricant return line3060.

FIG.27is a flowchart of a method3100of operating the lubrication system1600of the propulsion system2500ofFIG.24, according to the present disclosure. While reference is made to the lubrication system1600and the propulsion system2500, the method3100can be utilized for any of the lubrication systems detailed herein. The method3100can be performed by the one or more controllers2602(e.g., at least one of the first controller2602aor the second controller2602b). Further, while reference is made the second propulsor38b, the method3100can also be utilized for the first propulsor38a.

In step3105, the method3100includes operating the first propulsor38aand the second propulsor38b. In particular, the method3100includes operating the first turbine engine2910aand the second turbine engine2910b. For example, the first controller2602aoperates the first turbine engine2510a, and the second controller2602boperates the second turbine engine2510b. During the normal operation, the lubrication system1600supplies the lubricant to the gear assembly650and the engine bearings37, as detailed above.

In step3110, the method3100includes monitoring an operation status of the lubricant sensors1640of the second turbine engine2510b. Monitoring the operation status of the lubricant sensors1640of the second turbine engine2510bincludes at least one of comparing inputs from the dual channels of each lubricant sensor1640or comparing inputs from a plurality of the lubricant sensors1640. For example, comparing normal flow and normal pressure with a low lubricant level may indicate the lubricant level sensor1640ahas failed. In another example, normal flow and normal temperature with a low pressure may indicate the lubricant pressure sensor1640bhas failed.

In step3115, the method3100includes determining whether the operating status of the lubricant sensors1640of the second turbine engine2510bindicate a normal operation.

In step3120, the method3100includes operating the second propulsor38b(e.g., the second turbine engine2510b) at a normal power output for a particular operating mode (e.g., idle, taxi, takeoff, cruise, landing) if the operating status of the lubricant sensors1640of the second turbine engine2510bindicates the normal operation (step3115: YES). Operating the second propulsor38bat the normal power output includes at least one of a normal fuel flow rate, a normal power output from the second electric machine74b, or a normal pitch angle of the second propulsor blades40b. The normal operation includes operating the first turbine engine2510aand the second turbine engine2510bat the same power output level (or substantially equal power output levels). The method3100then proceeds back to step3105. If the operating status of the lubricant sensors1640is not normal (step3115: NO), the method3100proceeds to step3125.

In step3125, the method3100includes determining whether the operating status of the lubricant sensors1640of the second turbine engine2510bindicate a first failure mode. The first failure mode includes at least one of the lubricant temperature, the lubricant pressure, the lubricant level, or the lubricant flow rate to be greater than or less than a first threshold. For example, the first failure mode includes at least one of the lubricant temperature greater than a first lubricant temperature threshold, the lubricant pressure less than a first lubricant pressure threshold, the lubricant level less than a first lubricant level threshold, or the lubricant flow rate less than a first lubricant flow rate threshold.

In step3130, the method3100includes operating the second propulsor38b(e.g., the second turbine engine2510b) at a first reduced power output if the operating status indicates the first failure mode. The first reduced power output is less than the normal power output for the particular operating mode (e.g., idle, taxi, takeoff, cruise, landing). Operating the second propulsor38bat the first reduced power output includes at least one of a first reduced fuel flow rate, a first reduced power output from the second electric machine74b, or a first adjusted pitch angle of the second propulsor blades40b. A reduced power output level is such that the heat imparted to the lubricant (primarily by the gearbox gear mesh) is reduced in magnitude (e.g., a 10% reduction, a 20% reduction, or a 40% reduction in thrust or output power. The method3100then proceeds back to step3105. If the operating status of the lubricant sensors1640does not indicate the first failure mode (step3125: NO), the method3100proceeds to step3135.

In step3135, the method3100includes determining whether the operating status of the lubricant sensors1640of the second turbine engine2510bindicate a second failure mode. The second failure mode occurs when at least one of the lubricant temperature, the lubricant pressure, the lubricant level, or the lubricant flow rate is greater than or less than a second threshold. For example, the second failure mode includes at least one of the lubricant temperature is greater than a second lubricant temperature threshold, the lubricant pressure is less than a second lubricant pressure threshold, the lubricant level is less than a second lubricant level threshold, or the lubricant flow rate is less than a second lubricant flow rate threshold. The second lubricant temperature threshold is greater than the first temperature threshold. The second lubricant pressure threshold is less than the first lubricant pressure threshold. The second lubricant level threshold is less than the first lubricant level threshold. The second lubricant flow rate threshold is less than the first lubricant flow threshold.

In step3140, the method3100includes operating the second propulsor38b(e.g., the second turbine engine2510b) at a second reduced power output if the operating status indicates the second failure mode. The second reduced power output is less than the first reduced power output for the particular operating mode (e.g., idle, taxi, takeoff, cruise, landing). Operating the second propulsor38bat the second reduced power output includes at least one of a second reduced fuel flow rate, a second reduced power output from the second electric machine74b, or a second adjusted pitch angle of the second propulsor blades40b. The second reduced power output corresponds to a fail-safe level of operation at which the second reduced power output is a minimum power output of the second propulsor38bfor supplying the lubricant from the second lubrication system1600bto the first lubrication system1600a. The method3100then proceeds back to step3105. If the operating status of the lubricant sensors1640does not indicate the second failure mode (step3135: NO), the method3100proceeds to step3145.

In step3145, the method3100includes determining whether the operating status of the lubricant sensors1640of the second turbine engine2510bindicate a third failure mode. The third failure mode includes at least one of the lubricant temperature, the lubricant pressure, the lubricant level, or the lubricant flow rate to be greater than or less than a third threshold. For example, the third failure mode occurs when at least one of the lubricant temperature is greater than a third lubricant temperature threshold, the lubricant pressure is less than a third lubricant pressure threshold, the lubricant level is less than a third lubricant level threshold, or the lubricant flow rate is less than a third lubricant flow rate threshold. The third lubricant temperature threshold is greater than the second temperature threshold. The third lubricant pressure threshold is less than the second lubricant pressure threshold. The third lubricant level threshold is less than the second lubricant level threshold. The third lubricant flow rate threshold is less than the second lubricant flow threshold. In some embodiments, the third failure mode includes at least one of the lubricant pressure, the lubricant level, or the lubricant flow rate being zero. In some embodiments, the third failure mode is indicated by a chip detector fault in one or more of the lubricant sensors1640.

In such embodiments, the method3100includes setting a maintenance indicator flag in response to the chip detector fault, and outputting the maintenance indicator flag to a computing device. The computing device can include, for example, a display, a speaker, or the like, for outputting the maintenance indicator flag. For example, the maintenance indicator flag can include at least one of a visual alert on a display or an audio alert through a speaker. In some embodiments, the computing device is located on the aircraft (e.g., in the cockpit). In some embodiments, the computing device is located remote from the aircraft (e.g., on the ground).

In step3150, the method3100includes shutting down the second propulsor38b(e.g., the second turbine engine2510b). When the second propulsor38bis shut down, the power output is zero.

In step3155, the method3100includes feathering the second propulsor38bwhen the second propulsor38bis shut down. Feathering the second propulsor38bincludes adjusting the pitch of the second propulsor blades40bto a feather position (e.g., the chord of the second propulsor blades40bis approximately parallel to the flow of the volume of air50). Adjusting the pitch of the second propulsor blades40bto the feather position reduces drag on the second propulsor38bas the second propulsor38bcontinues to move through the air (e.g., during a flight). In some embodiments, the method3100includes applying a brake to the second propulsor38bwhen the second propulsor38bis shut down and the second propulsor blades40bare in the feather position. In this way, applying the brake prevents the second propulsor blades40bfrom rotating when the second propulsor38bis shut down. In some embodiments, applying the brake includes applying the brake on the second propulsor shaft45b. The method3100then proceeds back to step3105. If the operating status of the lubricant sensors1640does not indicate the third failure mode (step3145: NO), the method3100proceeds back to step3105.

FIG.28shows a computing system3200, according to the present disclosure. The computing system3200may carry out any of the methods or systems described herein. The controllers described previously herein may be according to the computing system3200. The computing system3200includes a general-purpose computing device, including a processing unit (CPU), or processor3220, and a system bus3210that couples various system components including a system memory3230such as a read-only memory (ROM)3240and a random-access memory (RAM)3250to the processor3220. The computing system3200can include a cache of high-speed memory connected directly with, in close proximity to, or integrated as part of the processor3220. The computing system3200copies data from the system memory3230or the storage device3260to the cache for quick access by the processor3220. In this way, the cache provides a performance boost that avoids the processor3220delays while waiting for data. These and other modules can control or be configured to control the processor3220to perform various actions. Other system memory3230may be available for use as well. The system memory3230can include multiple different types of memory with different performance characteristics. The disclosure may operate on a computing system3200with more than one processor3220or on a group or cluster of computing devices networked together to provide greater processing capability. The processor3220can include any general-purpose processor and a hardware module or software module, such as module 13262, module 23264, and module 33266stored in the storage device3260, configured to control the processor3220as well as a special-purpose processor where software instructions are incorporated into the actual processor design. The processor3220may essentially be a completely self-contained computing system, containing multiple cores or processors, a bus, memory controller, cache, etc. A multi-core processor may be symmetric or asymmetric.

The system bus3210may be any of several types of bus structures including a memory bus or memory controller, a peripheral bus, and a local bus using any of a variety of bus architectures. A basic input/output (BIOS) stored in the ROM3240or the like, may provide the basic routine that helps to transfer information between elements within the computing system3200, such as during start-up. The computing system3200further includes one or more storage devices3260such as a hard disk drive, a magnetic disk drive, an optical disk drive, a tape drive, or the like. The storage devices3260can include software modules3262,3264, and3266for controlling the processor3220. Other hardware or software modules are contemplated. The storage device3260is connected to the system bus3210by a drive interface. The drives and the associated computer-readable storage media provide nonvolatile storage of computer-readable instructions, data structures, program modules and other data for the computing system3200. In one aspect, a hardware module that performs a particular function includes the software component stored in a tangible computer-readable storage medium in connection with the necessary hardware components, such as the processor3220, the system bus3210, the output device3270, and so forth, to carry out the function. In another aspect, the computing system3200can use a processor and a computer-readable storage medium to store instructions which, when executed by a processor (e.g., one or more processors), cause the processor to perform a method or other specific actions. The basic components and appropriate variations are contemplated depending on the type of device, such as whether the computing system3200is a small, handheld computing device, a desktop computer, or a computer server.

Although certain embodiments described herein employs the storage device3260, other types of computer-readable media which can store data that are accessible by a computer, such as magnetic cassettes, flash memory cards, digital versatile disks, cartridges, random-access memories (RAMs)3250, and a read-only memory (ROM)3240, may also be used in the exemplary operating environment. Tangible computer-readable storage media, computer-readable storage devices, or computer-readable memory devices, expressly exclude media such as transitory waves, energy, carrier signals, electromagnetic waves, and signals per se.

To enable user interaction with the computing system3200, an input device3290represents any number of input mechanisms, such as a microphone for speech, a touch-sensitive screen for gesture or graphical input, keyboard, mouse, motion input, speech and so forth. An output device3270can also be one or more of a number of output mechanisms known to those of skill in the art. In some instances, multimodal systems enable a user to provide multiple types of input to communicate with the computing system3200. The communications interface3280generally governs and manages the user input and system output. There is no restriction on operating on any particular hardware arrangement and therefore the basic features here may easily be substituted for improved hardware or firmware arrangements as they are developed.

Accordingly, the present disclosure provides for propulsion systems that include hybrid electric turboprop engines. The turbine engine control systems herein can control the turbine engine and the remote propulsor based on a single throttle lever input. In this way, the propulsion systems of the present disclosure require only a single throttle lever to be actuated by the pilot to control both the turbine engine and the remote propulsor, thus, reducing pilot workload, allowing the pilot to better focus attention on other needs. The present disclosure also provides for a propulsion system that incorporates a gearbox assembly with a gearbox lubrication system that is separate from the engine lubrication system. The gearbox lubrication system can be powered by the electric machine or by the remote propulsor. In this way, the propulsion system can lubricate the gearbox assembly even when the turbine engine is shut down and the turbine shaft is no longer powering the pump of the engine lubrication system. The present disclosure also provides for a propulsion system having an electrically powered lubrication system that allows the turbo-engine to be shut down, while still providing lubrication the propulsor. The lubrication system sensors provide data to the remote engine as well as the local engine so as to allow for safe control of the local engine by the remote engine. Lastly, the present disclosure provides a plurality of lubricant sensors (e.g., temperature sensors, pressure sensors, level sensors, or flow rate sensors) associated with the lubrication system that enable fault monitoring by the remote turbine engine. The turbine engine control system identifies failure modes of the lubricant sensors and operates the remote turbine engine at a reduced power or shutting down the propulsor based on the failure modes.

An aircraft engine assembly comprising: a gas turbine engine having a high pressure compressor, a high pressure turbine, a high pressure shaft coupling the high pressure compressor with the high pressure turbine, a low pressure turbine, and a low pressure shaft coupled to the low pressure turbine, the high pressure turbine located forward of the high pressure compressor, and the low pressure turbine located on a forward end of the gas turbine engine; an intake channel of the gas turbine engine configured to receive an incoming flow of air and thereby form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction; and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.

The aircraft engine assembly of one or more of these clauses, which further includes a tail cone, and wherein the electric machine is located within the tail cone.

The aircraft engine assembly of one or more of these clauses, wherein the low pressure shaft is oriented to extend through the intake channel such that an axial portion of the low pressure shaft is surrounded by the incoming flow of air in the intake channel.

The aircraft engine assembly of one or more of these clauses, wherein the electric machine is positioned to receive an impingement of the intake flow of air.

The aircraft engine assembly of one or more of these clauses, wherein an offtake flow of air is extracted from the intake flow of air, the offtake flow of air forming a cooling flow of air routed to the electric machine.

The aircraft engine assembly of one or more of these clauses, wherein the low pressure turbine includes variable stator vanes, wherein the gas turbine engine includes an engine controller, and wherein the variable stator vanes of the low pressure turbine are controlled by the controller to change position when the electric machine changes from a first power level to a second power level.

The aircraft engine assembly of one or more of these clauses, which further includes a propeller located on a forward end of the gas turbine engine and coupled via the low pressure shaft with the low pressure turbine, and which further includes a gearbox coupled between the low pressure shaft and the propeller.

The aircraft engine assembly of one or more of these clauses, which further includes a propeller located on a forward end of the gas turbine engine and coupled via the low pressure shaft with the low pressure turbine, and wherein the propeller is variable pitch propeller, and wherein the electric machine is coaxial with the low pressure shaft.

The aircraft engine assembly of one or more of these clauses, which further includes an engine nacelle enclosing the gas turbine engine, the electric machine located in an enclosed space aft of the high pressure compressor.

The aircraft engine assembly of one or more of these clauses, wherein the engine nacelle includes a discharge opening sized to permit discharge of the portion of the intake flow of air.

The aircraft engine assembly of one or more of these clauses, wherein the electric machine is in a conductive heat exchange communication with the intake flow of air.

The aircraft engine assembly of one or more of these clauses, which further includes an engine nacelle enclosing the gas turbine engine, the electric machine located in an enclosed space aft of the high pressure compressor.

The aircraft engine assembly of one or more of these clauses, wherein the engine nacelle includes a discharge opening sized to permit discharge of the portion of the intake flow of air.

The aircraft engine assembly of one or more of these clauses, wherein the electric machine is in a conductive heat exchange communication with the intake flow of air.

The aircraft engine assembly of one or more of these clauses, wherein the intake channel includes a first portion that directs air into a first radial side of the low pressure shaft, wherein the intake channel includes a second portion configured as annular in shape, and wherein the first portion is upstream of the second portion, the second portion directing the intake flow of air in the first axial flow direction toward the high pressure compressor.

An aircraft powerplant comprising: a gas turbine engine having a high pressure compressor and a high pressure turbine, the gas turbine engine further having a high pressure shaft coupling the high pressure compressor with the high pressure turbine, the gas turbine engine also having a first axial flow direction from the high pressure compressor to the high pressure turbine; a propeller coupled to a low pressure turbine of the gas turbine engine using a low pressure shaft, the low pressure shaft located coaxial with the high pressure shaft, the propeller configured to receive a free stream flow of air oriented in a freestream direction and impart work upon the free stream flow of air, the propeller located on an upstream side of the freestream direction from the high pressure turbine; an intake channel defining an intake flow of air in fluid communication with the gas turbine engine, the intake channel configured to reverse the intake flow of air initially flowing in the freestream direction to the first axial flow direction of the gas turbine engine; and an electric machine coupled to the low pressure shaft and located on an opposite side of the gas turbine engine from the propeller, the electric machine positioned to be cooled by a portion of the intake flow of air defined by the intake channel.

The aircraft powerplant of one or more of these clauses, which further includes an engine nacelle enclosing the gas turbine engine, the electric machine located in an enclosed space aft of the high pressure compressor.

The aircraft powerplant of one or more of these clauses, wherein the engine nacelle includes a discharge opening sized to permit discharge of the portion of the intake flow of air.

The aircraft powerplant of one or more of these clauses, wherein the electric machine is in a conductive heat exchange communication with the intake flow of air.

The aircraft powerplant of one or more of these clauses, wherein the intake channel includes a first portion that directs air into a first radial side of the low pressure shaft, wherein the intake channel includes a second portion configured as annular in shape, and wherein the first portion is upstream of the second portion, the second portion directing air in the first axial flow direction toward the high pressure compressor.

The aircraft powerplant of one or more of these clauses, wherein the gas turbine engine includes an engine controller, and wherein the engine controller commands a change in fuel flow rate to a combustion section of the gas turbine engine when the electric machine changes from a first power level to a second power level.

The aircraft powerplant of one or more of these clauses, wherein the low pressure turbine is a free turbine.

The aircraft powerplant of one or more of these clauses, an engine exhaust configured to receive an exhaust flow in the first axial flow direction from the low pressure turbine and discharge the exhaust flow having a second axial direction component reverse of the first axial direction.

A method of cooling an electric machine comprising: operating a gas turbine engine having a high pressure compressor, a high pressure turbine, and a high pressure shaft, the high pressure compressor coupled to the high pressure turbine via the high pressure shaft, the gas turbine engine also including a low pressure turbine coupled via a low pressure shaft with a propeller; receiving an incoming flow of air in an incoming flow of air direction into an intake channel of the gas turbine engine, the incoming flow of air used in a combustion process of the gas turbine engine, the intake channel forming an intake channel flow path and defining an intake flow air; turning the intake flow of air using the intake channel from the incoming flow of air direction to an axial flow direction of the gas turbine engine; and cooling an electric machine located aft of the low pressure turbine using the intake flow of air flowing through the intake channel flow path.

The method of cooling an electric machine of one or more of these clauses, wherein the cooling includes impingement cooling the electric machine using the intake flow of air.

The method of cooling an electric machine of one or more of these clauses, wherein turning the flow includes changing shape of the flow from a non-annular shape to an annular shape.

The method of cooling an electric machine of one or more of these clauses, wherein the electric machine is located in a tail cone located aft of the gas turbine engine.

An aircraft engine assembly comprising: a gas turbine engine having a high pressure compressor, a high pressure turbine, a high pressure shaft coupling the high pressure compressor with the high pressure turbine, a low pressure turbine, and a low pressure shaft coupled to the low pressure turbine, the high pressure turbine located forward of the high pressure compressor, and the low pressure turbine located on a forward end of the gas turbine engine; a first propeller located on a forward end of the gas turbine engine and coupled via the low pressure shaft with the low pressure turbine; a second propeller located remote from the gas turbine engine; an intake channel of the gas turbine engine configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction; an electric machine coupled with the low pressure shaft and located on a side of the high pressure compressor opposite of the high pressure turbine and proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated; and a turbine engine control system comprising: a single throttle lever; and a controller that: receives an input from the single throttle lever; and controls the gas turbine engine and the second propeller based on the input from the single throttle lever.

The aircraft engine assembly of any preceding clause, wherein the controller controls at least one of a rotational speed of the turbo-engine, a rotational speed of the first propeller, a pitch of a plurality of first propeller blades of the first propeller, or a torque of the first propeller.

The aircraft engine assembly of any preceding clause, wherein the controller controls at least one of a rotational speed of the second propeller, a pitch of a plurality of second propeller blades of the second propeller, or a torque of the second propeller.

The aircraft engine assembly of any preceding clause, wherein the controller receives a position input from the single throttle lever, converts the position input into at least one of a turbo-engine setting, a first propeller setting, or a second propeller setting, and controls the second propeller and at least one of the turbo-engine or the first propeller based on the at least one of the turbo-engine setting, the first propeller setting, or the second propeller setting.

The aircraft engine assembly of any preceding clause, further comprising an aircraft power supply that powers the second propeller.

The aircraft engine assembly of any preceding clause, wherein the gas turbine engine is a first gas turbine engine and the turbo-engine is a first turbo-engine, and further comprising a second turbine engine having a second turbo-engine, and the second propeller being drivingly coupled to the second turbo-engine.

The aircraft engine assembly of any preceding clause, wherein the controller controls the second turbine engine based on the input from the single throttle lever when the second turbine engine is in a standby operating mode or a failure mode.

The aircraft engine assembly of any preceding clause, wherein the electric machine is drivingly coupled to the first propeller, the electric machine powering the first propeller when the gas turbine engine is shut down.

The aircraft engine assembly of any preceding clause, wherein the electric machine is a first electric machine, and further comprising a second electric machine that is drivingly coupled to the second propeller, the second electric machine powering the second propeller.

The aircraft engine assembly of any preceding clause, further comprising a power conduit that provides electrical communication from the first electric machine to the second electric machine, the first electric machine powering the second electric machine through the power conduit.

The aircraft engine assembly of any preceding clause, further comprising a gearbox coupled between the low pressure shaft and the first propeller.

The aircraft engine assembly of any preceding clause, further comprising a gearbox lubrication system.

The aircraft engine assembly of any preceding clause, wherein the gearbox lubrication system further comprises a gearbox lubricant reservoir, a pump, and a gearbox lubricant supply line.

The aircraft engine assembly of any preceding clause, wherein the gearbox lubrication system further comprises a heat exchanger.

The aircraft engine assembly of any preceding clause, wherein the pump is in fluid communication with the low pressure shaft.

The aircraft engine assembly of any preceding clause, wherein the gearbox lubrication system further comprises a lubricant sensor.

The aircraft engine assembly of any preceding clause, wherein the controller is in communication with the lubricant sensor.

The aircraft engine assembly of any preceding clause, wherein the lubricant sensor is a first lubricant sensor and the controller is a first controller, the first lubricant sensor and the first controller associated with the first propeller, and further comprising a second lubricant sensor and a second controller associated with the second propeller.

The aircraft engine assembly of any preceding clause, wherein the lubricant sensor includes a lubricant level sensor and a lubricant pressure sensor.

The aircraft engine assembly of any preceding clause, wherein the gearbox lubrication system further comprises one or more valves controllable by the controller.