Patent Description:
Various types and configurations of geartrains for an aircraft propulsion system are known in the art for an aircraft. While these known aircraft propulsion system geartrains have various benefits, there is still room in the art for improvement.

<CIT> relates to gearboxes through which an aircraft engine shaft can drive a load.

<CIT> relates to electrically controlled vertical takeoff and landing aircraft systems.

According to the invention, an assembly is provided for an aircraft as claimed in claim <NUM>. Various optional embodiments of the invention are provided as claimed in the dependent claims.

<FIG> schematically illustrates a propulsion system <NUM> for an aircraft. The aircraft may be an airplane, a helicopter, a drone (e.g., an unmanned aerial vehicle (UAV)), a spacecraft or any other manned or unmanned aerial vehicle. This aircraft may be configured as a vertical take-off and landing (VTOL) aircraft or a short take-off and vertical landing (STOVL) aircraft. The aircraft propulsion system <NUM> of <FIG>, for example, is configured to generate power for first direction propulsion (e.g., propulsive thrust) during a first mode of operation and to generate power for second direction propulsion (e.g., propulsive lift) during a second mode of operation, where the first direction is different than (e.g., angularly offset from) the second direction. The first mode may be a horizontal flight mode (e.g., a forward flight mode) where the first direction propulsion is substantially horizontal propulsive thrust; e.g., within five degrees (<NUM>°), ten degrees (<NUM>°), etc. of a horizontal axis. The second mode may be a vertical flight and/or hover mode where the second direction propulsion is substantially vertical propulsive lift; e.g., within five degrees (<NUM>°), ten degrees (<NUM>°), etc. of a vertical axis. The aircraft propulsion system <NUM>, of course, may also be configured to generate both the first direction propulsion (e.g., horizontal propulsion) and the second direction propulsion (e.g., vertical propulsion) during a third mode (e.g., a transition mode) of operation.

The aircraft propulsion system <NUM> of <FIG> includes one or more bladed propulsor rotors such as, for example, at least one bladed first propulsor rotor <NUM> and at least one bladed second propulsor rotor <NUM>. The aircraft propulsion system <NUM> of <FIG> also includes a gas turbine engine core <NUM> configured to rotatably drive the one or more propulsor rotors - the first propulsor rotor <NUM> and/or the second propulsor rotor <NUM>.

The first propulsor rotor <NUM> may be configured as a ducted rotor such as a fan rotor. The first propulsor rotor <NUM> of <FIG> is rotatable about a first rotor axis <NUM>. This first rotor axis <NUM> is an axial centerline of the first propulsor rotor <NUM> and may be horizontal when the aircraft is on ground and/or during level aircraft flight. The first propulsor rotor <NUM> includes at least a first rotor disk <NUM> and a plurality of first rotor blades <NUM> (one visible in <FIG>); e.g., fan blades. The first rotor blades <NUM> are distributed circumferentially around the first rotor disk <NUM> in an annular array. Each of the first rotor blades <NUM> is connected to and projects radially (relative to the first rotor axis <NUM>) out from the first rotor disk <NUM>.

The second propulsor rotor <NUM> may be configured as an open rotor such as a propeller rotor or a helicopter (e.g., main) rotor. Of course, in other embodiments, the second propulsor rotor <NUM> may alternatively be configured as a ducted rotor such as a fan rotor; e.g., see dashed line duct. The second propulsor rotor <NUM> of <FIG> is rotatable about a second rotor axis <NUM>. This second rotor axis <NUM> is an axial centerline of the second propulsor rotor <NUM> and may be vertical when the aircraft is on the ground and/or during level aircraft flight. The second rotor axis <NUM> is angularly offset from the first rotor axis <NUM> by an included angle <NUM>; e.g., an acute angle or a right angle. This included angle <NUM> may be between sixty degrees (<NUM>°) and ninety degrees (<NUM>°); however, the present disclosure is not limited to such an exemplary relationship. The second propulsor rotor <NUM> includes at least a second rotor disk <NUM> and a plurality of second rotor blades <NUM>; e.g., open rotor blades. The second rotor blades <NUM> are distributed circumferentially around the second rotor disk <NUM> in an annular array. Each of the second rotor blades <NUM> is connected to and projects radially (relative to the second rotor axis <NUM>) out from the second rotor disk <NUM>.

The engine core <NUM> extends axially along a core axis <NUM> between a forward, upstream airflow inlet <NUM> and an aft, downstream exhaust <NUM>. The core axis <NUM> may be an axial centerline of the engine core <NUM> and may be horizontal when the aircraft is on the ground and/or during level aircraft flight. This core axis <NUM> may be parallel (e.g., coaxial) with the first rotor axis <NUM> and, thus, angularly offset from the second rotor axis <NUM>. The engine core <NUM> of <FIG> includes a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. The turbine section <NUM> of <FIG> includes a high pressure turbine (HPT) section 48A and a low pressure turbine (LPT) section 48B (also sometimes referred to as a power turbine section).

The engine sections <NUM>-48B are arranged sequentially along the core axis <NUM> within an engine housing <NUM>. This engine housing <NUM> includes an inner case <NUM> (e.g., a core case) and an outer case <NUM> (e.g., a fan case). The inner case <NUM> may house one or more of the engine sections <NUM>-48B; e.g., the engine core <NUM>. The outer case <NUM> may house the first propulsor rotor <NUM>. The outer case <NUM> of <FIG> also axially overlaps and extends circumferentially about (e.g., completely around) the inner case <NUM> thereby at least partially forming a (e.g., annular) bypass flowpath <NUM> radially between the inner case <NUM> and the outer case <NUM>.

Each of the engine sections <NUM>, 48A and 48B includes a bladed rotor <NUM>-<NUM> within that respective engine section <NUM>, 48A, 48B. Each of these bladed rotors <NUM>-<NUM> includes a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s).

The compressor rotor <NUM> is connected to the HPT rotor <NUM> through a high speed shaft <NUM>. At least (or only) these engine components <NUM>, <NUM> and <NUM> collectively form a high speed rotating structure <NUM>. This high speed rotating structure <NUM> is rotatable about the core axis <NUM>. The LPT rotor <NUM> is connected to a low speed shaft <NUM>. At least (or only) these engine components <NUM> and <NUM> collectively form a low speed rotating structure <NUM>. This low speed rotating structure <NUM> is rotatable about the core axis <NUM>. The low speed rotating structure <NUM> and, more particularly, its low speed shaft <NUM> may project axially through a bore of the high speed rotating structure <NUM> and its high speed shaft <NUM>.

The aircraft propulsion system <NUM> of <FIG> includes a powertrain <NUM> that couples the low speed rotating structure <NUM> to the first propulsor rotor <NUM> and that couples the low speed rotating structure <NUM> to the second propulsor rotor <NUM>. The powertrain <NUM> of <FIG> includes a geartrain <NUM>, a transmission <NUM> and a gearing <NUM>; e.g., bevel gearing. The powertrain <NUM> of <FIG> also includes one or more shafts <NUM>-<NUM> and/or other torque transmission devices for coupling the geartrain <NUM> to the first propulsor rotor <NUM> and the second propulsor rotor <NUM>.

An input to the geartrain <NUM> is coupled to the low speed rotating structure <NUM> and its low speed shaft <NUM>, where the low speed rotating structure <NUM> forms a power input for the geartrain <NUM>. A first output from the geartrain <NUM> is coupled to the first propulsor rotor <NUM> through the first propulsor shaft <NUM>, where the first propulsor rotor <NUM> forms a first power output (e.g., load) for the geartrain <NUM>. A second output from the geartrain <NUM> is coupled to the second propulsor rotor <NUM> through the powertrain elements <NUM>, <NUM>, <NUM>, <NUM> and <NUM>, where the second propulsor rotor <NUM> forms a second power output (e.g., load) for the geartrain <NUM>.

An output of the transmission <NUM> is connected to the gearing <NUM> through the transmission output shaft <NUM>. This transmission <NUM> may be configured to selectively couple (e.g., transfer mechanical power between) the geartrain output shaft <NUM> and the transmission output shaft <NUM>. During the first mode of operation, for example, the transmission <NUM> may be configured to decouple the geartrain output shaft <NUM> from the transmission output shaft <NUM>, thereby decoupling the low speed rotating structure <NUM> from the second propulsor rotor <NUM>. During the second mode of operation (and the third mode of operation), the transmission <NUM> may be configured to couple the geartrain output shaft <NUM> with the transmission output shaft <NUM>, thereby coupling the low speed rotating structure <NUM> with the second propulsor rotor <NUM>. The transmission <NUM> may be configured as a clutched or clutchless transmission.

An output of the gearing <NUM> is connected to the second propulsor rotor <NUM> through the second propulsor shaft <NUM>. This gearing <NUM> provides a coupling between the transmission output shaft <NUM> rotating about the axis <NUM>, <NUM> and the second propulsor shaft <NUM> rotating about the second rotor axis <NUM>. The gearing <NUM> may also provide a speed change mechanism between the transmission output shaft <NUM> and the second propulsor shaft <NUM>. The gearing <NUM>, however, may alternatively provide a <NUM>:<NUM> rotational coupling between the transmission output shaft <NUM> and the second propulsor shaft <NUM> such that these shafts <NUM> and <NUM> rotate at a common (e.g., the same) speed. Furthermore, in some embodiments, the gearing <NUM> and the transmission output shaft <NUM> may be omitted where the functionality of the gearing <NUM> is integrated into the transmission <NUM>. In still other embodiments, the transmission <NUM> may be omitted where decoupling of the second propulsor rotor <NUM> is not required and/or where an optional additional speed change between the second output of the geartrain <NUM> and the second propulsor rotor <NUM> is not required.

During operation of the aircraft propulsion system <NUM>, air enters the engine core <NUM> through the airflow inlet <NUM>. This air is directed into a (e.g., annular) core flowpath <NUM> which extends sequentially through the compressor section <NUM>, the combustor section <NUM>, the HPT section 48A and the LPT section 48B to the exhaust <NUM>. The air within this core flowpath <NUM> may be referred to as core air.

The core air is compressed by the compressor rotor <NUM> and directed into a (e.g., annular) combustion chamber <NUM> of a (e.g., annular) combustor <NUM> in the combustor section <NUM>. Fuel is injected into the combustion chamber <NUM> through one or more fuel injectors <NUM> (one visible in <FIG>) and mixed with the compressed core air to provide a fuel-air mixture. This fuel-air mixture is ignited and combustion products thereof flow through and sequentially cause the HPT rotor <NUM> and the LPT rotor <NUM> to rotate. The rotation of the HPT rotor <NUM> drives rotation of the high speed rotating structure <NUM> and its compressor rotor <NUM>. The rotation of the LPT rotor <NUM> drives rotation of the low speed rotating structure <NUM>. The rotation of the low speed rotating structure <NUM> drives rotation of the first propulsor rotor <NUM> through the geartrain <NUM> during a select mode or modes of operation; e.g., the first and the third modes of operation. The rotation of the low speed rotating structure <NUM> drives rotation of the second propulsor rotor <NUM> through the geartrain <NUM> during a select mode or modes of operation; e.g., the second and the third modes of operation. During the first mode of operation, the transmission <NUM> may decouple the low speed rotating structure <NUM> from the second propulsor rotor <NUM> such that the low speed rotating structure <NUM> does not drive rotation of the second propulsor rotor <NUM>. The second propulsor rotor <NUM> may thereby be stationary (or windmill) during the first mode of operation.

During the first and the third modes of operation, the rotation of the first propulsor rotor <NUM> propels bypass air (separate from the core air) through the aircraft propulsion system <NUM> and its bypass flowpath <NUM> to provide the first direction propulsion; e.g., the forward, horizontal thrust. During the second and the third modes of operation, the rotation of the second propulsor rotor <NUM> propels additional air (separate from the core air and the bypass air) to provide the second direction propulsion; e.g., vertical lift. The aircraft may thereby takeoff, land and/or otherwise hover during the second and the third modes of operation, and the aircraft may fly forward or otherwise move during the first and the third modes of operation.

<FIG> and <FIG> illustrate aspects of the geartrain <NUM> in further detail. The geartrain <NUM> of <FIG> and <FIG> includes a sun gear <NUM>, a ring gear <NUM>, a plurality of intermediate gears <NUM> and a carrier <NUM>. The sun gear <NUM> is rotatable about a rotational axis <NUM> of the geartrain <NUM>, which rotational axis <NUM> may be parallel (e.g., coaxial) with the axis <NUM>, <NUM>. The ring gear <NUM> circumscribes the sun gear <NUM>, and the ring gear <NUM> is rotatable about the axis <NUM>, <NUM>, <NUM>. Each of the intermediate gears <NUM> is disposed (e.g., radially) between and meshed with the sun gear <NUM> and the ring gear <NUM>. Each of the intermediate gears <NUM> is rotatably mounted to the carrier <NUM>. The carrier <NUM> is rotatable about the axis <NUM>, <NUM>, <NUM>. The first propulsor rotor <NUM> is coupled to and is configured to be rotatably driven by the carrier <NUM>. The second propulsor rotor <NUM> is coupled to and is configured to be rotatably driven by the sun gear <NUM>. The low speed rotating structure <NUM> and its LPT rotor <NUM> are coupled to and is configured to rotatably drive the ring gear <NUM>. With this arrangement, the geartrain <NUM> may provide a propulsor rotor (e.g., fan) to LPT rotor speed ratio capable of configuring the first propulsor rotor <NUM> with a relatively high pressure ratio (e.g., fan pressure ratio (FPR)) while rotating the low speed rotating structure <NUM> and its LPT rotor <NUM> at relatively fast rotational speeds.

Referring to <FIG>, the aircraft propulsion system <NUM> may include a propulsor rotation control system <NUM>. This propulsor rotation control system <NUM> may include one or more brakes 104A and 104B (generally referred to as "<NUM>") and/or one or more lock devices 106A and 106B (generally referred to as "<NUM>").

The first brake 104A and/or the first lock device 106A may be located at location <NUM>, or another suitable location. The first brake 104A is configured to brake (e.g., slow and/or stop) rotation of the first propulsor rotor <NUM> about the axis <NUM>, <NUM>, <NUM>. The first lock device 106A is configured to lock (e.g., fix, prevent) rotation of the first propulsor rotor <NUM> about the axis <NUM>, <NUM>, <NUM>, for example, following the braking of the first propulsor rotor <NUM> to a zero rotational speed about the axis <NUM>, <NUM>, <NUM> using the first brake 104A. When the first propulsor rotor <NUM> is rotationally fixed (e.g., during the second mode of operation of <FIG>), the geartrain <NUM> may transfer (e.g., all, minus losses in the powertrain <NUM>) the power output from the low speed rotating structure <NUM> and its LPT rotor <NUM> to the second propulsor rotor <NUM> and the powertrain element(s) therebetween.

The second brake 104B and/or the second lock device 106B may be located at location <NUM>, or another suitable location. The second brake 104B is configured to brake (e.g., slow and/or stop) rotation of the second propulsor rotor <NUM> about the axis <NUM>. The second lock device 106B is configured to lock (e.g., fix, prevent) rotation of the second propulsor rotor <NUM> about the axis <NUM>, for example, following the braking of the second propulsor rotor <NUM> to a zero rotational speed about the axis <NUM>. When the second propulsor rotor <NUM> is rotationally fixed (e.g., during the first mode of operation of <FIG>), the geartrain <NUM> may transfer (e.g., all, minus losses in the powertrain <NUM>) the power output from the low speed rotating structure <NUM> and its LPT rotor <NUM> to the first propulsor rotor <NUM> and the powertrain element(s) therebetween.

To enter the third mode of operation from the first mode of operation, the second lock device 106B may be disengaged and/or the second brake 104B may be released. The second propulsor rotor <NUM> may thereby begin to rotate along with the already rotating first propulsor rotor <NUM>. Similarly, to enter the third mode of operation from the second mode of operation, the first lock device 106A may be disengaged and/or the first brake 104A may be released. The first propulsor rotor <NUM> may thereby begin to rotate along with the already rotating second propulsor rotor <NUM>. When both of the propulsor rotors <NUM> and <NUM> are rotating / free to rotate (e.g., during the third mode of operation of <FIG>), the geartrain <NUM> may transfer (e.g., all, minus losses in the powertrain <NUM>) the power output from the low speed rotating structure <NUM> and its LPT rotor <NUM> to (I) the first propulsor rotor <NUM> and the powertrain element(s) therebetween and (II) the second propulsor rotor <NUM> and the powertrain element(s) therebetween.

Referring to <FIG>, the first brake 104A and/or the second brake 104B may each be configured as or otherwise include a disk brake <NUM>. The disk brake <NUM> of <FIG> includes a brake rotor <NUM> and one or more brake pads <NUM>. The brake rotor <NUM> is configured rotatable with the respective propulsor rotor <NUM>, <NUM>. The brake rotor <NUM>, for example, may be connected to and rotatable with the respective shaft <NUM>, <NUM>, or another rotating element (directly or indirectly) rotatable with the respective propulsor rotor <NUM>, <NUM>. The brake pads <NUM> are anchored to a stationary structure <NUM>, which may be part of the engine housing <NUM> and/or an airframe of the aircraft (see <FIG>). The brake pads <NUM> may be actuated by one or more brake actuators <NUM> (e.g., hydraulic brake actuators) to move the brake pads <NUM> from an open position to a closed position. In the open position, the brake pads <NUM> are spaced from and do not engage (e.g., contact) the brake rotor <NUM> (see position of <FIG>). In the closed position, the brake pads <NUM> engage (e.g., contact) and clamp onto (e.g., squeeze) the brake rotor <NUM>. Frictional rubbing between the brake pads <NUM> and the brake rotor <NUM> is operable to brake rotation of the brake rotor <NUM> and, thus, the respective shaft <NUM>, <NUM> (or other rotating element) connected thereto. The first and the second brakes <NUM> of the present disclosure, however, are not limited to such an exemplary disk brake configuration. Furthermore, it is contemplated the first and/or the second brake <NUM> may alternatively be configured as another type of brake such as, for example, a drum brake.

Referring to <FIG>, the first lock device 106A and/or the second lock device 106B may each be configured as a splined lock device <NUM>; e.g., a splined coupling. The lock device <NUM> of <FIG>, for example, includes an inner lock element <NUM> (e.g., a splined shaft), an outer lock element <NUM> (e.g., a splined sleeve) and an actuator <NUM>. The inner lock element <NUM> is rotatable about the axis <NUM>, <NUM>, <NUM>. The outer lock element <NUM> is rotationally fixed about the axis <NUM>, <NUM>, <NUM>. However, the actuator <NUM> is configured to move (e.g., axially translate) the outer lock element <NUM> along the axis <NUM>, <NUM>, <NUM> and the inner lock element <NUM> between an unlocked position (see dashed line in <FIG>) and a locked position (see solid line in <FIG>; see also <FIG>). At the unlocked position, inner splines <NUM> of the outer lock element <NUM> are disengaged (e.g., spaced) from outer splines <NUM> of the inner lock element <NUM>. At the locked position, the inner splines <NUM> of the outer lock element <NUM> are engaged (e.g., meshed) with the outer splines <NUM> of the inner lock element <NUM> (see also <FIG>). With this arrangement, when the lock device <NUM> is unlocked and its outer lock element <NUM> is in the unlocked position, the inner lock element <NUM> may rotate (e.g., freely, unencumbered by the outer lock element <NUM>) about the axis <NUM>, <NUM>, <NUM>. However, when the lock device <NUM> is locked and its outer lock element <NUM> is in the locked position of <FIG>, the outer lock element <NUM> is meshed with the inner lock element <NUM> and prevents rotation of the inner lock element <NUM> about the axis <NUM>, <NUM>, <NUM>.

Referring to <FIG> and <FIG>, the inner lock element <NUM> of the first lock device 106A may be configured as part of or may be attached (directly or indirectly) to the first propulsor shaft <NUM>, or any other element rotatable therewith. The inner lock element <NUM> of the second lock device 106B may be configured as part of or may be attached (directly or indirectly) to the geartrain output shaft <NUM>, or any other element rotatable therewith. While the inner lock element <NUM> of <FIG> and <FIG> is described as the rotating element and the outer lock element <NUM> is described as the rotationally fixed element, the operation of these elements may be switched in other embodiments. In particular, the inner lock element <NUM> may alternatively be configured as the rotationally fixed element and axially translatable by the actuator <NUM>, and the outer lock element <NUM> may be configured as the rotating element. Furthermore, various other types of rotational lock devices are known in the art, and the present disclosure is not limited to any particular ones thereof.

In some embodiments, referring to <FIG>, the low speed rotating structure <NUM> may be configured without a compressor rotor. In other embodiments, referring to <FIG> (see also <FIG>), the low speed rotating structure <NUM> may include a low pressure compressor (LPC) rotor <NUM>' arranged within a low pressure compressor (LPC) section 46A of the compressor section <NUM>. In such embodiments, the compressor rotor <NUM> may be a high pressure compressor (HPC) rotor within a high pressure compressor (HPC) section 46B of the compressor section <NUM>.

The engine core <NUM> (e.g., see <FIG>) may have various configurations other than those described above. The engine core <NUM>, for example, may be configured with a single spool, with two spools (e.g., see <FIG>), or with more than two spools. The engine core <NUM> may be configured with one or more axial flow compressor sections, one or more radial flow compressor sections, one or more axial flow turbine sections and/or one or more radial flow turbine sections. The engine core <NUM> may be configured with any type or configuration of annular, tubular (e.g., CAN), axial flow and/or reverser flow combustor. The present disclosure therefore is not limited to any particular types or configurations of gas turbine engine cores. Furthermore, it is contemplated the engine core <NUM> of the present disclosure may drive more than the two propulsor rotors <NUM> and <NUM>, or a single one of the propulsor rotors <NUM> and <NUM> and/or one or more other mechanical loads; e.g., electric machines, electric generators, electric motors, etc. The aircraft propulsion system <NUM>, for example, may include two or more of the first propulsor rotors <NUM> and/or two or more of the second propulsor rotors <NUM>. For example, the aircraft propulsion system <NUM> of <FIG> includes multiple second propulsor rotors <NUM> rotatably driven by the low speed rotating structure <NUM>. These second propulsor rotors <NUM> may rotate about a common axis. Alternatively, each second propulsor rotor <NUM> may rotate about a discrete axis where, for example, the second propulsor rotors <NUM> are laterally spaced from one another and coupled to the low speed rotating structure <NUM> through a power splitting geartrain <NUM>.

Claim 1:
An assembly for an aircraft, comprising:
a geartrain (<NUM>) comprising a sun gear (<NUM>), a ring gear (<NUM>), a plurality of intermediate gears (<NUM>) and a carrier (<NUM>), the ring gear (<NUM>) circumscribing the sun gear (<NUM>) and rotatable about an axis (<NUM>), each of the plurality of intermediate gears (<NUM>) between and meshed with the sun gear (<NUM>) and the ring gear (<NUM>), each of the plurality of intermediate gears (<NUM>) rotatably mounted to the carrier (<NUM>), and the carrier (<NUM>) rotatable about the axis (<NUM>);
a first propulsor rotor (<NUM>) coupled to the carrier (<NUM>); and
a rotating structure (<NUM>) coupled to the ring gear (<NUM>), the rotating structure (<NUM>) comprising a turbine rotor (<NUM>), and the rotating structure (<NUM>) configured to drive rotation of the first propulsor rotor (<NUM>) through the geartrain (<NUM>),
wherein the sun gear (<NUM>) is rotatable about the axis (<NUM>), characterised in that the assembly further comprises:
a second propulsor rotor (<NUM>) coupled to the sun gear (<NUM>); and
the rotating structure (<NUM>) configured to drive rotation of the second propulsor rotor (<NUM>) through the geartrain (<NUM>).