AIRCRAFT PROPULSION SYSTEM

An aircraft propulsion system includes at least first and second propulsor arrangements. Each propulsor arrangement includes an internal combustion engine configured to mechanically drive a plurality of propulsors. At least one of the propulsors of each propulsor arrangement is driveable via a shaft and at least one reduction gearbox. Each internal combustion engine is coupled to a motor generator arrangement configured to drive at least one of the mechanically driven propulsors when in a motor mode, and to generate electrical power when in a generating mode. The system further includes an electrical interconnector configured to transfer electrical power between the first and second propulsor arrangements.

The present disclosure concerns an aircraft propulsion system.

Powered aircraft such as civilian airliners are typically propelled by propulsors in the form of either open rotor propellers or ducted fans, which are typically mechanically driven by either gas turbine engines or reciprocating piston engines. Where a ducted fan is driven by a gas turbine engine, such an arrangement is known as a “turbofan” or “bypass jet”, whereas where an open propeller is driven by a gas turbine engine, such an arrangement is known as a “turboprop”, “open rotor gas turbine” or “propfan”.

In general, there is a trend towards fewer, larger engines in civilian airliners. A typical arrangement in which one relatively large engine is provided on each wing is relatively efficient, whilst providing redundancy. This is because the thrust Specific Fuel Consumption (SFC) of gas turbine engines generally decreases with engine size, so it is generally more efficient to have fewer, larger engines. Furthermore, the drag associated with additional engine nacelles is eliminated. Meanwhile, it is a regulatory requirement to provide at least two engines in many cases, such that the aircraft can continue to operate in the event of one engine being inoperative (OEI).

There is also a continuing trend within the aviation industry to provide turbofans and turboprops having a higher bypass ratio, i.e. having a larger proportion of the air mass flow through the fan/propeller disc, rather than through the engine core. In general, larger bypass ratios result in lower SFC. However, as the fan increases in size (i.e. diameter) with high bypass ratio engines, several problems are encountered. These include for example, high fan tip speeds, excessively heavy engines nacelles and fan containment systems (in the case of turbofans), and issues with ground clearance.

One proposal to maintain a subsonic fan tip speed in a large diameter fan is to provide a reduction gearbox between the turbine and fan of the gas turbine engine, such that the fan rotates at a lower speed than the turbine. Such an arrangement is provided for example in the Pratt and Whitney PW1000G™. Such arrangements are also common in turboprops. However, as the fan/propeller diameter is increased further, the rotational speed of the fan must be reduced in order to maintain a subsonic fan/propeller tip speed (which is necessary to avoid the tips becoming supersonic, and so creating large amounts of noise), which in turn results in the gearbox having to handle a large amount of torque for a given power rating, due to the increased reduction ratio. Consequently, the gearbox may be relatively heavy, and it may not be possible to scale such an arrangement up to higher thrusts or bypass ratios.

An alternative solution is to drive a plurality of ducted or unducted fans using a common gas turbine engine core, thereby permitting each fan to have a smaller diameter, while maintaining a high bypass ratio for the propulsion system as a whole. One such example is disclosed in U.S. Pat. No. 8,402,740, in which bevel gears are used to power two non-coaxial fans from a single gas turbine engine core shaft. U.S. Pat. No. 8,015,796 describes an arrangement in which a layshaft gearbox is used to transfer power to the non-coaxial fans. Another alternative solution is to transfer power from one or more gas turbines to a plurality of remotely sited fans or propellers via electrical generators and an electrical transmission network. Such arrangements are described for example in Gohardani, A. S. ‘A synergistic glance at the prospects of distributed propulsion technology and the electric aircraft concept for future unmanned air vehicles and commercial/military aviation’, Progress in Aerospace Sciences, Volume 57, February 2013. However, such arrangements introduce inefficiencies in view of the requirement to convert mechanical power to electrical power, and back again. These disadvantages may be partly overcome using superconducting generators, motors and cables. However, these technologies are relatively immature, and can be expected to add additional weight and cost.

Each of these solutions generally increases the weight of the powerplant, and so may result in a net reduction (or only very slight net increase) in overall aircraft level efficiency, at significantly increased cost. Furthermore, there must be provision made for OEI operation, and so at least two separate powerplants (i.e. gas turbines or reciprocating engines) are required. In the case of an electric distributed propulsion aircraft, provision must be made for the failure of either a gas turbine engine, an electrically driven propulsor, or a generator, leading to potentially extensive redundancy, and so high cost. Even so, where the engines are installed on the wings of the aircraft, a large amount of rudder deflection will be required in the event of the loss of one gas turbine engine in view of the resultant asymmetric thrust. This may restrict the design of the aircraft to particular takeoff and landing speeds (necessitated by the Vmca, i.e. the minimum airspeed at which the aircraft is controllable with one engine out), and may also necessitate a relatively large rudder and possibly also relatively large roll control surfaces, which may in turn result in a relatively large amount of drag.

The present invention provides an aircraft propulsion system and an aircraft which seeks to address one or more of the above problems.

According to a first aspect of the invention there is provided an aircraft propulsion system comprising:

at least first and second propulsor arrangements, each propulsor arrangement comprising an internal combustion engine, a plurality of propulsors, a shaft, a reduction gearbox arrangement and a motor generator arrangement, the internal combustion engine being configured to drive the plurality of propulsors via the shaft and the reduction gearbox arrangement, each internal combustion engine being coupled to the motor generator arrangement, the motor generator arrangement being configured to drive at least one of the mechanically driven propulsors when in a motor mode, and to generate electrical power when in a generating mode; and
an electrical interconnector configured to transfer electrical power between the first and second propulsor arrangements.

Advantageously, the arrangement provides an aircraft propulsion system which provides efficient operation, while ensuring that power can be transmitted to at least one of the propulsors of a propulsor arrangement having an inoperative gas turbine engine, due to the motor generator arrangement and the electrical interconnector. Consequently, benefits at an aircraft level can be achieved in view of reduced yaw requirements in the event of the failure of a single gas turbine engine. Furthermore, since a large number of propulsors are provided, the remaining propulsors would only have to provide a moderately increased thrust in the event of a propulsor failure, thereby reducing the nominal rating of each propulsor, and leading to weight and cost advantages.

At least one of the propulsors may comprise one of a ducted fan and an open rotor propeller.

Each of the first and second propulsor arrangements may be mounted to a respective wing.

One or more of the propulsors may be located having an inlet located upstream of a leading edge of an aircraft wing. One or more of the propulsors may be located such that a wing flap is located in a slipstream of the respective propulsor. The invention is thought to be particularly advantageous in such an arrangement, as the slipstream produced by the propellers increases the effectiveness of the flaps. Consequently, the wing can be made smaller, thereby reducing net aircraft drag. However, in such arrangements, a lift imbalance would occur where propulsors on one wing were to be inoperative. In view of the electrical interconnector and motor generators coupled to propulsors, this disadvantage is reduced or eliminated.

The propulsion system may comprise at least one propulsor located so as to ingest a boundary layer airflow in use. The propulsion system may comprise at least one propulsor having an inlet located rearwardly of a trailing edge of the wing. Advantageously, the propulsors ingest a boundary layer airflow in use, thereby reaccelerating boundary airflow, and so improving the propulsive efficiency of the aircraft.

The system may comprise a tip propulsor comprising a propulsor mounted to a wing tip, having an inlet adjacent the wing tip. The tip propulsor may be electrically driven, and may be located having an inlet located downstream of a leading edge of the wing tip, and may be located having an inlet located downstream of a trailing edge of the wing tip. The system may comprise a tip propulsor controller configured to control thrust generated by the tip propulsor in accordance with a yaw demand. Advantageously, the tip propulsor can be used to provide yaw control, thereby reducing the impact of an OEI yaw imbalance, thereby in turn allowing a further reduction in vertical stabiliser/rudder surface area. Thus the disclosed system may have a reduced weight and aerodynamic drag compared to prior systems. In view of the positioning of the tip propulsor inlet adjacent the wing tip, wing tip vortices can be reduced. It has been found that, in many instance, the maximum power required by the gas turbine engines in a high bypass ratio propulsion system is defined by the power required for “second segment” climb, i.e. climb to a high altitude after takeoff. Under such conditions, in which the wings must operate at high lift coefficients, vortices are shed from the wingtip. The vortices represent a significant contributor to drag in second segment climb, and so reducing these vortices can be expected to reduce the power requirement during this phase of flight, and so reduce the power requirement of the gas turbine engines. Consequently, the gas turbine engines can be made lighter in view of the reduced power requirements, resulting in large efficiencies beyond the direct efficiencies produced by reducing drag associated with wing tip vortices.

The system may comprise a boundary layer ingesting electrically driven propulsor mounted at a rearward end of an aircraft fuselage, having an inlet downstream of a trailing edge of the aircraft fuselage.

The one or more shafts may be configured to disconnect in the event of a failure of one of a coupled component. For example, the one or more shafts may comprise a clutch or frangible link. Advantageously, in the event of a failure of one of a coupled component such as a gearbox, a propulsor or a motor generator, the remaining coupled components can continue to operate.

The reduction gearbox arrangement may comprise a bevel gearbox configured to transfer shaft power from a gas turbine engine driving shaft to a first shaft having an axis of rotation generally coaxial to the gas turbine engine driving shaft, and to a second shaft having an axis of rotation generally normal to the gas turbine engine driving shaft.

At least one motor generator may be coupled to the second shaft.

The propulsion system may comprise one or more further bevel gearboxes configured to transfer shaft power from the second shaft to a propulsor driving output shaft having an axis of rotation generally normal to the axis of rotation of the second shaft.

According to a second aspect of the present disclosure there is provided an aircraft comprising a propulsion arrangement in accordance with the first aspect of the disclosure.

The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects of the invention may be applied mutatis mutandis to any other aspect of the invention.

Referring toFIG. 1, a twin-spooled, gas turbine engine is generally indicated at10. The engine10comprises a core engine11having, in axial flow series, an air intake12, a low pressure compressor14, a high-pressure compressor15, combustion equipment16, a high-pressure turbine17, a low-pressure turbine18and a core exhaust nozzle20. A nacelle21generally surrounds the core engine11and defines the intake12, nozzle20and a core exhaust duct22. High and low pressure shafts8,9couple the high and low pressure compressors14,15and turbines17,18respectively. The low pressure shaft24extends forward of the core engine11to drive a load.

The gas turbine engine10works in a conventional manner so that air entering the intake12is accelerated and compressed by the intermediate pressure compressor14and directed into the high-pressure compressor15where further compression takes place. The compressed air exhausted from the high-pressure compressor15is directed into the combustion equipment16where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high pressure and low pressure turbines17,18before being exhausted through the nozzle20to provide some propulsive thrust.

The gas turbine engine10is part of a propulsion system100of an aircraft30, shown in more detail inFIG. 2.

Referring toFIG. 2, the aircraft30comprises a fuselage32which defines a longitudinal axis, and provides internal space for passengers and cargo. Attached to the fuselage is a pair of wings34and an empennage36comprising vertical and horizontal surfaces.

The propulsion system comprises a pair of propulsor arrangements100a,100b, each of which is mounted to a respective wing34. One of the propulsor arrangements is shown schematically in more detail inFIG. 3.

Each propulsor arrangement100a,100bcomprises a two spool gas turbine engine10mounted to a respective wing34within a nacelle21, and a plurality of mechanically driven propulsors130a-c. The low pressure shaft9of the gas turbine engine10drives a reduction gearbox104, shown in further detail inFIG. 4.

Referring toFIG. 4, the reduction gearbox104is contained within a housing105. The low pressure shaft9is the input shaft for the reduction gearbox104, and drives an input bevel gear108. The input bevel gear108is configured to transfer power to three output shafts106a,106b,106cvia respective intermediate bevel gears110,112and final bevel gear114respectively. The axes of rotation of gears108,110,112,114and shafts9,106a-care arranged generally orthogonally, such that the input bevel gear108drives intermediate bevel gears110,114, and intermediate bevel gears110,114in turn drive final bevel gear112. Gears108,110,114have diameters and teeth numbers such that a reduction ratio of approximately 4:1 or higher is provided between the input shaft9and output shafts106a,106b,106c. It will be understood though that different reduction gear ratios could be employed. In order to provide such a reduction on all three output shafts106a-cwhile maintaining a 90° angle between the input shaft9and output shafts106a,106c, gear112and shaft106bcould be offset to the other gears108,110,114in a direction normal to their rotational axes, for example in a vertical direction (i.e. into or out of the page in the diagram shown inFIG. 4).

Alternatively, the gearbox could provide a different reduction ratio for different output shafts. For example, the gears108,110,114could have the same number of teeth, with gear112having a greater number of teeth, such that the output shafts106a,106crotate at the same speed as the input shaft9, with output shaft106brotating at a lower rotational rate than the input shaft9. Advantageously, such an arrangement can provide relatively high reduction ratios of 4 or greater, since an epicyclic gearbox is not required. Consequently, the gas turbine engine spools can rotate at relatively high speeds (which results in high turbine tip speeds without requiring large diameter turbine discs), while the propulsors130a-ccan operate at low speeds (which results in efficient propulsion in view of the low tip speeds). In contrast, the reduction ratio of epicyclic gearbox is generally limited to around 4:1. High reduction gear ratios will also reduce the torque requirements of the shafts and driven equipment.

Shafts106aand106bextend in generally opposite directions to one another, and have axes of rotation generally normally to the input shaft9, and extent in a generally spanwise direction. Consequently, the gearbox104splits power from the low pressure input shaft9into two outputs having a different axis of rotation to the input shaft9, and a lower rotational speed. Meanwhile, the output shaft106bis arranged to rotate generally coaxially with the input shaft24.

Each propulsor arrangement100further comprises a pair of motor generators116. The motor generators116comprise a stator118having one or more magnetic poles arranged around a rotor120. The rotor120is coupled to a respective shaft106a,106c, such that the rotor120and shaft106a,160cco-rotate. Consequently, where the motor generators116are in a generating mode (i.e. where the respective motor generator116is being driven by the gas turbine engine10via respective shaft106a,106b), rotation of the shaft106agenerates electrical power, whereas where the motor generator is in a motor mode (i.e. being provided with electrical power), the electrical power provided to the stator118causes the motor to drive the respective shaft106a,106b. The respective shaft106a,106c, rotor120and stator118are arranged concentrically, and generally have a high aspect ratio, such that the motor generators116extend for a substantial portion of the overall length of the shaft106a,106c. Consequently, the motor116is relatively compact, having a relatively small diameter. As such, the motor116can generally be located within the wing34. In general, the motor/generator is located close to the leading edge of the wing, where the wing is thickest. Each motor generator116comprises an AC motor generator, such as a synchronous or asynchronous motor.

The shafts106a,106care coupled to a further respective bevel gearbox122at a distal end thereof, and thereby provide an input shaft of the respective bevel gearbox122. The bevel gearboxes122comprise a pair of bevel gears124,126arranged generally orthogonally, such that an output shaft128of the bevel gearbox has an axis of rotation generally normal to the axis of rotation of the input shaft106a,106c. In the described embodiment, the bevel gearbox translates the axis of rotation, but does not provide a reduction gear. Such an arrangement, in conjunction with the gearbox104, ensures that the propellers130a,130crotate in opposite directions to propeller130b. Such an arrangement reduces the P-factor, and reduces aerodynamic interference between the propellers, thereby increasing propulsive efficiency.

The output shaft128of each bevel gearbox122is coupled to a propulsor in the form of an open rotor propeller130a,130c. Each propeller130a,130chas at least one propeller blade132attached by a hub133, and is configured to provide thrust when rotated. Similarly, shaft106bdirectly drives a further propeller130b. Consequently, the gas turbine engine10drives the propellers130a,130b,130cvia the reduction gearbox104, and also via bevel gearbox122in the case of propellers130a,130c. In view of the relatively large number of propellers (i.e. three in the described embodiment), the propulsion system100can have a large effective bypass ratio whilst having relatively small diameter propellers130a-c. Consequently, the propellers can rotate at relatively high speed without encountering supersonic tip speeds. As a result, the reduction ratio provided by the reduction gearbox arrangement can be relatively low, such that the engine shaft24can run at a relatively high speed (which results in a relatively efficient turbine), while the output shafts run at only a slightly lower speed. As a consequence, relatively small, low weight gearboxes can be provided.

Each propulsor arrangement further comprises at least one shaft disconnection arrangement in the form of a plurality of frangible connections146. Each frangible connection is configured to uncouple a respective shaft106a,106cfrom the remaining shafts and gearboxes in the event of a failure of one of the gearboxes, shafts or mechanically driven propulsors130a-c. For instance, the frangible connections146may be configured to physically break where a maximum load is exceeded. Consequently, failure of one mechanically driven component will not propagate to other components, thereby providing additional redundancy.

The propulsor arrangements100are electrically interconnected by an interconnector140. The interconnector140is an electrical connector which electrically couples the motor generators116of the left propulsor arrangement100a, with the motor generators116of the right propulsor arrangement100b. Consequently, in the event of a failure of the gas turbine engine10of one of the propulsor arrangements, power can be transferred from one propulsor arrangement100a,100bto the other electrically. For example, where the left propulsor arrangement100agas turbine engine10fails in flight, the motor generators116of the right propulsor arrangement100bwould be operated in a generator mode, while the motor generators116of the left propulsor arrangement100awould be operated in a motor mode. Consequently, the propulsors130a-cof the left propulsor arrangement100awould continue to operate in OEI conditions, when one gas turbine engine10has failed. In such circumstances, a load on the shafts106of the right propulsor arrangement100bwould be produced by the motor generators116operating in generator mode. Consequently, it may be desirable for the propulsors103to comprise variable pitch rotors132, such that the aerodynamic load on the propulsors130can be modified in order to accommodate the increased shaft load. In other words, the pitch may need to be reduced in the event of OEI operation, so that the propulsors can continue to operate at high rotational speed. In some cases, the power transferred between propulsor arrangements100a,100bunder OEI conditions may be less than 50%—i.e. the propulsor arrangement100a,100bhaving the operative gas turbine engine10may provide a greater proportion of the power than the arrangement100a,100bhaving the inoperative gas turbine engine10. Since the interconnector may experience some resistive losses (estimated at perhaps 5% of transmitted power), the propulsive efficiency of the system as a whole can be increased by transmitting less than 50% of the power through the interconnector140. The resultant thrust imbalance can be accommodated by utilising tip propulsors134, or using the rudder.

Referring once more toFIG. 3, the mechanically driven propulsors130a,130cof each propulsion system100a,100bare located upstream of trailing edge flaps144of each wing34. Consequently, the flaps144are located in the slipstream of the propulsors130a-c. Such an arrangement is known in the art as “externally blown flaps” or “powered lift”. As a result, the wings34can be operated at a greater coefficient of lift compared to where only a single, relatively small diameter propulsor is used. Consequently, the wing area can be made smaller, while still providing sufficient lift for takeoff at acceptably low speeds. As a result, total airframe drag is reduced. Alternatively, steeper descents can be made. Meanwhile, the redundancy provided by the interconnector140and motor generators116ensures that the propulsors130a-cof both propulsor arrangements100a,100bcontinue to operate with one gas turbine engine10inoperative, thereby preventing the situation where lift is lost on one wing due to one of the propulsors no longer providing thrust. On the other hand, the gas turbine engine inlets12are also located within the slipstream of the propulsor130b, though this need not be the case.

The propulsion system100optionally further comprises at least one tip propulsor134mounted to a tip138of each wing34, having the propeller blades132located aft of a trailing edge of each wing34. Each propulsor134is driven by an electric motor138, which is provided with electrical power from the motor generators116via the electrical interconnector140. Consequently, the tip propulsors134are located at a point where a wingtip vortex would normally be generated. By rotating the propellers in a clockwise direction as viewed from downstream of the propulsor134on the port wing34, and in an opposite direction on the starboard wing, the wingtip vortex can be at least partly cancelled, thereby reducing the wake vortex. This arrangement would be expected to reduce losses due to tip vortices, and also allow closer spacing between aircraft in congested airspace.

As a further advantage, each tip propulsor134is located a large distance from the centre of mass of the aircraft30, and so the thrust generated by each propulsor134may provide a significant yawing moment. By controlling each propulsor134via a controller136in accordance with a yaw requirement (as determined either manually by the pilot, or in accordance with an OEI schedule), the size of the rudder can be reduced, thereby further reducing drag and weight. In some cases, the rudder, and possibly the vertical tail surface, can be eliminated entirely. On the other hand, since the propulsors134are driven by an electric motor provided with electrical power from the propulsor arrangement (rather than a mechanical shaft), the power can be transmitted through the wing without limiting the flexibility of the wing (which would require increasing the stiffness, and therefore the weight of the wing34), or by requiring flexible couplings, which have increased complexity. In view of the relatively low power requirements of the tip propulsors134, the electrical interconnector140can be relatively light. Since the tip propulsors134are powered by electrical power provided by the interconnector140, electrical power can continue to be provided during OEI operation. Consequently, these propulsors will be particularly advantageous in cancelling any adverse yaw in the event of OEI operation.

A further electrically driven propulsor is provided in the form of a boundary layer ingesting propulsor142mounted with the propeller blades132being located within the boundary layer at the aft end of the fuselage32, downstream of the empennage36. Again, electrical power is provided to the propulsor142from the gas turbine engines10via the interconnector140. The boundary layer ingesting propulsor142ingests the boundary layer generated by the fuselage32, thereby increasing the propulsive efficiency of the propulsion system100. Again, in view of the relatively low power requirements of the propulsor142, a relatively low weight electrical interconnector144can be utilised. Furthermore, in view of the presence of the electrical interconnection140between the first and second propulsion arrangements100a,100b, at least part of the weight of the electrical interconnector is already accounted for in the design. Consequently, the additional weight of electrical cabling to provide a boundary layer ingesting propulsor at an aft end of the aircraft in addition to the interconnection between the gas turbine engines10is relatively small.

For example, the disclosed propulsion system may be suitable for different types of aircraft, such as blended wing aircraft, in which the fuselage provides lift, such that there is no distinctive separation between the fuselage and wings. Alternatively, the aircraft could have a canard configuration, in which the horizontal tail surfaces are omitted, and replaced by a canard located at a forward end of the fuselage.

The propulsors could be of different types, such as for example ducted fans. The propulsors could be located on different parts of the aircraft. For example, the mechanically driven propulsors could be located at a trailing edge of the aircraft. Such an arrangement would ensure that airflow over the wings is undisturbed by propeller wash, while also ingesting an amount of boundary layer air, thereby increasing the propulsive efficiency of the propulsors. However, such an arrangement may not be compatible with trailing edge flaps, which may necessitate higher landing speeds. Either or both of the tip propulsors or the rear fuselage mounted boundary layer ingesting propulsors could be omitted. Alternatively, further mechanically and/or electrically driven propulsors could be provided.

Other types of motor generator arrangements could be provided. For example, the motor generator arrangement could comprise separate motor and electrical generator units coupled to the same shaft. Different numbers of motor generators could be provided. For example, a single motor generator could be provided for each propulsor arrangement, and could for example be mounted directly to a shaft of the respective gas turbine engine, i.e. not through a reduction gearbox.

Other gearbox arrangements could be provided. For example, power could be mechanically transmitted from the gas turbine engine to the propulsors via a layshaft or planetary gearbox. Instead of a single reduction gearbox, multiple reduction gearboxes could be provided, such that, for instance, the bevel gearbox104could have a reduction ratio for only the coaxial output shaft, while rpm reduction could be provided by bevel gearboxes122. Alternatively, the gearboxes104,122could provide no ratio reduction, with ratio reduction being provided by separate gearboxes provided for each propulsor. The system may also include a step-up gearbox, to increase the rotational speed of shafts carrying the motor/generators, to increase the efficiency and/or reduce the size of the motor gearboxes, with the bevel gearboxes providing a reduction gearbox to reduce the rotational speed of the propellers. Similarly, the motor/generators could be located on a different shaft, such as the output shaft of the bevel gearbox, which may result in more efficient/more aerodynamic packaging of the motor/generator.

The electrical interconnectors could comprise superconducting cables. The motor generators could comprise superconducting electrical machines.

Though the gas turbine engines are shown as having intake located in the slipstream of the propulsors, the gas turbine engines could be arranged to ingest freestream air.