Patent Publication Number: US-2023160357-A1

Title: Aircraft propulsion system

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority pursuant to 35 U.S.C. 119(a) to United Kingdom Application No. 2116848.9, filed Nov. 23, 2021, which application is incorporated herein by reference in its entirety. 
     FIELD 
     The present disclosure concerns a propulsion system for an aircraft and an aircraft comprising the propulsion system. 
     BACKGROUND 
     Electric Propulsion system for aircraft have been proposed, in which one or more electric motors is employed to drive one or more propulsors. Hybrid propulsion systems are also known, in which a gas turbine engine or other internal combustion engine is used to drive the propulsors in concert with the electric motors. “Parallel hybrid” systems can be distinguished from so-called “series hybrid” systems, in that in a parallel hybrid system, a mechanical connection is provided between the internal combustion engine and at least one propulsor, with at least one electric motor driving either the same propulsor as that driven by the internal combustion engine, or a further propulsor. In a “series hybrid”, the internal combustion engine is not mechanically coupled to any propulsor, but merely drives a generator to supply electric power to one or more motors. 
     SUMMARY 
     According to a first aspect there is provided an aircraft propulsion system comprising; a propulsor; a motor; and a reduction gearbox coupled to a prime mover at an input side, and the propulsor at an output side, the reduction gearbox being configured to provide a reduction ratio between the input and output; wherein the reduction gearbox is configured such that the input side rotates in an opposite direction to the output side, and the prime mover, motor, propulsor and reduction gearbox are configured such that gyroscopic forces of the propulsor, motor, prime mover and reduction gearbox generated during aircraft manoeuvres and/or propulsion system failures are substantially cancelled. 
     Advantageously, by counter-rotating the motor and propulsor, and by arranging the motor, gearbox and propulsor mass distributions and reduction gearing in a particular manner, gyroscopic forces can be substantially eliminated during aircraft manoeuvres. This can reduce bending forces in use, which may reduce fatigue and increase aircraft life. Furthermore, the structural mass that would normally be required to oppose these forces is reduced, thereby reducing the weight of the propulsion system. 
     The prime mover may comprise the motor and/or an internal combustion engine such as a gas turbine engine. The motor may be provided on the input or output side of the reduction gearbox. 
     The reduction gearbox may comprise an epicyclic gearbox in the form of a star gearbox comprising a sun gear defining the input side, one or more planet gears which mesh with the sun gear, a statically mounted planet carrier configured to rotatably mount the planet gears, and a ring gear which meshes with the planet gears, the ring gear being configured to rotate to define the output side. Advantageously, input and output side rotational directions are reversed, and a reduction ratio is provided, in a single unit. 
     The propulsion system may comprise one or more further electrical motors each coupled to a respective planet gear. Advantageously, gyroscopic forces can be eliminated or reduced, while design freedom is increased, since the further motor can be used to supplement power, while also adding to the rotational inertia of the gearbox. Consequently, the designer can accommodate a wider range of motor sizes and reduction ratios. 
     The propulsion system may comprise a further electric motor coupled to the output side of the gearbox. Advantageously, gyroscopic forces can be eliminated or reduced, while design freedom is increased, since the further motor can be used to supplement power, while also adding to the rotational inertia of the output side of the fan. Consequently, the designer can accommodate a wider range of motor sizes and reduction ratios compared to where a single motor is employed. 
     The gearbox may comprise a reduction ratio of between 1.5:1 and 3.5:1. 
     The electric motor may comprise one of an axial flux motor and a radial flux motor. The electric motor may comprise a rotor radially inward of a stator, or a rotor radially outward of a stator. 
     The propulsor may comprise one of an open rotor propeller and a ducted fan. 
     The motor, propulsor and reduction gearbox may be configured such that, at a given input and output rotational speed, the total angular momentum of clockwise rotating components is within 50% of the sum of total angular momentum of anti-clockwise rotating components. Consequently, the angular momentum of clockwise and anti-clockwise rotating components is substantially cancelled, thereby reducing or substantially eliminating gyroscopic forces. 
     Preferably, at a given input and output rotational speed, the total angular momentum of clockwise rotating components is within 20%, preferably within 10%, and more preferably within 5% of the sum of total angular momentum of anti-clockwise rotating components. 
     According to a second aspect of the invention there is provided a method of designing an aircraft propulsion system comprising defining a propulsor; defining a motor; and defining a reduction gearbox coupled to the motor at an input side, and the propulsor at an output side, the reduction gearbox being configured to provide a reduction ratio between the input and output, the reduction gearbox being configured such that the input side rotates in an opposite direction to the output side; and designing the motor, propulsor and reduction gearbox such that gyroscopic forces of the propulsor, motor and reduction gearbox generated during aircraft manoeuvres and/or propulsion system failures are substantially cancelled. 
     The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore, except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein. 
    
    
     
       BRIEF DESCRIPTION 
       Embodiments will now be described by way of example only, with reference to the Figures, in which: 
         FIG.  1    is a plan view of a first aircraft comprising a propulsion system; 
         FIG.  2    is a schematic diagram of a propulsion system for the aircraft of  FIG.  1   ; 
         FIG.  3    is a schematic diagram of an alternative propulsion system for the aircraft of  FIG.  1   ; 
         FIG.  4    is a schematic diagram of a further alternative propulsion system for the aircraft of  FIG.  1   ; 
         FIG.  5    is a schematic side view of an electric motor suitable for the propulsion system shown in any of  FIGS.  2  to  4   ; 
         FIG.  6    is a schematic side view of an alternative electric motor suitable for the propulsion system shown in any of  FIGS.  2  to  4   ; 
         FIG.  7    is a schematic side view of a further alternative electric motor suitable for the propulsion system shown in any of  FIGS.  2  to  4   ; 
         FIG.  8    is a schematic diagram of a parallel hybrid propulsion system for the aircraft of  FIG.  1   ; 
         FIG.  9    is a schematic diagram of an alternative parallel hybrid propulsion system for the aircraft of  FIG.  1   ; and 
         FIG.  10    is a schematic diagram of a further alternative parallel hybrid propulsion system for the aircraft of  FIG.  1   . 
     
    
    
     DETAILED DESCRIPTION 
     With reference to  FIG.  1   , an aircraft  1  is shown. The aircraft is of conventional configuration, having a fuselage  2 , wings  3 , tail  4  and a pair of propulsion systems  5 . One of the propulsion systems  5  is shown in figure detail in  FIG.  2   . 
       FIG.  2    shows the propulsion system  5  schematically. The propulsion system  5  includes a propulsor  12 . In this embodiment, the propulsor  12  comprises an open rotor propeller, but could alternatively comprise a ducted fan. The propulsor  5  is powered by an electrical energy source. In this embodiment, the energy source comprises a battery  56 , which is in turn charged by a generator  58  driven by a gas turbine engine  60 . Consequently, the aircraft can be considered to be a “series hybrid” aircraft. 
     The propulsor  12  is configured to provide thrust for the aircraft  1 , and is coupled to a fan shaft  14  by bearings (not shown) which provide for rotation in use. The fan shaft  14  is in turn coupled to an output shaft  18  of a reduction gearbox  16 . 
     The reduction gearbox  16  is in the form of an epicyclic star gearbox, comprising a sun gear  20 , one or more planet gears  22  which mesh with the sun gear, a planet carrier  24  configured to rotatably mount the planet gears  22 , and a ring gear  26  which meshes with the planet gears  22 . The sun gear  20  is provided at a radially inner position, the ring gear  26  is provided at a radially outer position, and the planet gears  22  are provided therebetween. 
     The reduction gearbox  16  is configured as a star gearbox, in which each of the sun gear  20 , planet gears  22  and ring gear  26  are mounted for rotation about their respective axes, while the planet carrier  24  is statically mounted, such that the planet gears  22  do not precess around the sun gear  20 . Suitable bearing arrangements (not shown) are provided, to allow for rotation of each component, while providing the necessary support. The sun gear  20  is utilised as the input. The ring gear  26  is utilised as the output, and so is coupled to the fan  12  via the output shaft  18 . A reduction ratio is provided between the input and output, which is equal to the number of ring gear teeth divided by the number of sun gear teeth. In one example, the reduction ratio is approximately 3:1, and is typically between 1.5 and 3.5:1. 
     The propulsion system  5  further comprises a prime mover in the form of an electric motor  28  which is configured to drive the propulsor  12  via the reduction gearbox  16 . The motor  28  is of a conventional type, such as an induction or permanent magnet electric machine. In the present embodiment, the motor  28  is coupled to the fan  12  via the sun gear  20  and an input shaft  30 . Suitable electric motor types are shown in  FIGS.  5  to  7   . In the example shown in  FIG.  5   , the electric motor comprises an “in-runner” radial flux permanent magnet electric motor comprising a permanent magnet rotor  32  which rotates about an axis X. The rotor  32  is surrounded by a stator  34 , provided radially outwardly of the rotor  32 , separated by an air gap  36 . The stator  34  comprises electrical windings (not shown), which can be energised to produce a rotating magnetic field. This rotating magnetic field interacts with a magnetic field produced by the permanent magnets of the rotor  32 , to cause rotation. Since the magnetic flux between the rotor  32  and stator  34  crosses the air gap  36  in a radial direction. 
     The electric motor  28  is coupled to an energy storage device (not shown) in the form of one or more of a chemical battery, fuel cell, and capacitor, which provides the electric motor  28  with electrical power during operation. In some cases, multiple energy storages systems, which may be of different types (chemical battery, fuel cell etc) may be provided for each propulsion system  5 . In other cases, a common energy storage device may be provided for multiple propulsion systems. 
     As shown in  FIG.  2   , the propulsion system comprises first and second struts  36 ,  38 . The first strut  36  mounts the gearbox planet carrier  24  to the aircraft pylon or other appropriate static structure, while the second strut  38  mounts the motor  28  to the aircraft static structure. 
       FIG.  6    shows another suitable electric motor  128 . The motor in this case is an “out-runner” radial flux electric motor. The motor  128  is similar to the motor  28 , but the rotor  132  and stator  134  are reversed, such that the stator  134  is located radially inward of the rotor  132 . 
     Similarly,  FIG.  7    shows a further suitable electric motor  228 . In this case, the electric motor comprises an “axial flux” electric motor comprises a rotor  232  and stator  234 . The rotor and stator are spaced in an axial direction, with an air gap  236  extending therebetween. Since the flux crosses the air-gap in the axial direction, the motor  228  is described as an axial flux machine. 
     In each case, the rotational speed of the rotor  32 ,  132 ,  232  of the motor is higher than the rotational speed of the fan  12 . Electric machines are typically limited by their torque density (i.e. their maximum torque divided by their weight) rather than their power density (i.e. their maximum power divided by their weight). It is therefore desirable to operate electric machines at high rotational speed, to thereby maximise the mechanical power generated by the motor  28 ,  128 ,  228 . On the other hand, it is desirable to rotate the fan  12  at a relatively low rotational speed, in order to minimise fan tip speed, to reduce aerodynamic losses. Consequently, the reduction gearbox  16  enables the motor  28  to turn at a higher rate than the fan  12 . However, the gearbox  16  leads to an increased overall propulsion system  5  weight, and also represents a transmission loss (often in the range of approximately 1% of motor power). Consequently, it may be difficult to determine whether a direct drive or a reduction gearbox is preferable. The inventors have determined however that the reduction gearbox leads to overall system weight reduction, if certain criteria apply (as set out below). 
     It will be noted that the propulsion system  5  comprises several rotating components. These include the fan  12 , fan shaft  14 , output shaft  18 , ring gear  26 , planet gears  22 , sun gear  20 , motor shaft  30  and motor rotor  32 . The bearings also represent rotating components, but can be ignored in this example in view of their relatively low mass. Each of these components will produce a gyroscopic moment, i.e. a force which resists turning in a direction normal to the rotational axis. In the event of an aircraft manoeuvre (such as yaw, pitch or roll), this gyroscopic moment will resist the turning moment imposed by the aircraft  1  on the propulsion system  5 . This will in turn induce a bending force on structural components, in particular, struts  36 ,  38 . Consequently, the more rapid the manoeuvre and the larger the rotational inertia of the rotating components, the larger the force will be. 
     These forces come into play in two situations in particular. At take-off, the aircraft rotates rapidly (i.e. pivots in the pitch axis). At this condition, the motor  28  and fan  12  will be running at high speed, and so the forces imparted on the struts  36 ,  38 , and other structures will be high. 
     Separately, rotational inertia forces also come into play in the event of a sudden stop of a rotating component. For example, in the event of a fan shaft break, or a gearbox seizure, there may be a sudden acceleration or deceleration of one or more rotating components. This sudden acceleration will result in large torque moments applied to the static structure  36 ,  38 , which must again be reacted. 
     Consequently, the static components  36 ,  38  must react loads in multiple directions, including twisting and pivoting directions. These loads may be high in view of the large rotational inertia in the propulsion system  5 . 
     The inventors have however found that these gyroscopic and inertial moments can be minimised, by careful design of the motor  28 , fan  12  and gearbox  16 . 
     In the present disclosure, the gearbox  16  is arranged such that the input side (comprising the sun gear  20 , motor shaft  30  and motor rotor  32 ) rotate in a first common direction (which could be clockwise or anticlockwise as desired). This is shown by the downward pointing arrow in  FIG.  2   . Since the gearbox  16  is a star gearbox configuration, in which the planet carrier  24  is held static, and the ring gear  26  is used as an output, the output side of the gearbox  16  rotates in an opposite direction to the input side. In this case, the rotating components of the output side of the gearbox  16  includes the fan  12 , fan shaft  14 , gearbox output shaft  18  and ring gear  26 . The planet gears  22  also rotate in the same direction as the output side. 
     Each of the rotating components  12 ,  14 ,  18 ,  26 ,  20 ,  22 ,  30 ,  32  has a corresponding mass moment of inertia (rotational moment of inertia), i.e. the second moment of mass with respect to the distance from the axis. When rotating, the rotating components each have an angular momentum, and the propulsion system as a whole will have a total angular momentum equal to the vector sum of the individual angular momenta of the rotating components  12 ,  14 ,  18 ,  26 ,  20 ,  22 ,  30 ,  32 . 
     As will be understood, the gyroscopic force imparted on the static structure  36 ,  38  is related to the angular momentum of the rotating components. By minimising the total angular momentum, preferably reducing this close to zero, the gyroscopic force is minimised. This will allow the static structure  36 ,  38  to be weaker and therefore lighter. 
     A first means for reducing the total angular momentum would be to reduce the mass of the components  12 ,  14 ,  18 ,  26 ,  20 ,  22 ,  30 ,  32 , or design the components such that mass is carried at a smaller diameter. However, this severely restricts design freedom. For instance, it is desirable to have a large fan  12  diameter, to increase propulsive efficiency. Similarly, small diameter motors must rotate more quickly for a given power output, which may increase the required gear reduction to achieve a desired fan rotational speed. 
     The inventors have found that gyroscopic forces can instead be cancelled by arranging some components to rotate in a first direction (say, clockwise), and others in a second direction (say, anti-clockwise). 
     In the present disclosure, the sun gear  20 , motor shaft  30  and motor rotor  32  rotate in a clockwise direction, and the fan  12 , fan shaft  14 , gearbox output shaft  18 , planet gears  22  and ring gear  26  rotate in an anti-clockwise direction. By arranging the masses, relative rotational speeds (by altering the gear reduction ratio) and radial mass distributions of these components, the gyroscopic forces can be substantially cancelled at a given rotational speed. Since the speed relationship between the input and outputs sides is typically fixed (in view of a reduction gearbox  16  having a fixed reduction ratio), gyroscopic forces can be cancelled for substantially any rotational speed. 
     In one example, the rotational inertia of the fan  12 , fan shaft  14 , gearbox output shaft  18 , planet gears  22  and ring gear  26  is equal to twice the rotational inertia of the sun gear  20 , motor shaft  30  and motor rotor  32 , and the gearbox  16  is arranged to have a reduction ratio of 2:1. Consequently, since the rotational speed of the clockwise rotating components is twice the rotational speed of the anti-clockwise rotating components, while the rotational inertia of the clockwise rotating components is half the rotational inertia of the ant-clockwise rotating components, the angular momentum of the input and output sides is equal and opposite at any given rotational speed, and so the gyroscopic forces are entirely cancelled. 
     In practice, the designer may adjust the mass distributions of the various components, and the reduction ratio, to achieve a desired combination of features, while substantially cancelling, or at least minimising, gyroscopic forces. 
     In practice, it may not be necessary to entirely cancel gyroscopic forces, and may not even be desirable. For instance, arranging the system to entirely cancel gyroscopic forces may lead to an undesirable compromise (for example, a non-optimal reduction ratio, non-optimal fan diameter, gearbox diameter, non-optimal motor speed, etc). Consequently, the inventors have found that it is desirable to reduce gyroscopic forces to less than 20% of what the forces would be if all the components rotated in the same direction. In other words, the rotating components are configured such that, at a given rotational speed, the sum of the angular momentum of the clockwise rotating components of the propulsor, motor and reduction gearboxes is within 20% of the sum of the of the angular momentum of each anti-clockwise rotating component of the propulsor, motor and reduction gearboxes. In other words, the vector sum of the angular momentum of the rotating components is compared to the sum of the absolute values (ignoring the different sign due to the opposite direction of rotation) of the angular momenta of the rotating components at a given speed. If the vector sum is less than 20% of the sum of the absolute values, then the gyroscopic forces can be said to be substantially cancelled. 
     In some cases, the gyroscopic forces may need to be cancelled to a greater degree, say between 1% and 10% of the sum of the absolute values of the rotating components at a given speed, or between 1 and 5% of the sum of the absolute values of the rotating components at a given speed. 
       FIG.  3    shows an alternative propulsion system  305 . The system  305  is similar to the system  5 , and comprises a propulsive fan  312  driven by a fan shaft  314 , which is in turn driven by a star reduction gearbox  316 , which is in turn driven by a first electric motor  328  via an input shaft  330 . The propulsion system is mounted by a static mounting system  336 ,  338 . Each of the above components are similar to those of the system  5 , and so are not described in further detail. 
     The propulsion system  305  further comprises a second motor  340 , which is coupled to the fan  312  via the fan shaft  314 , and so rotates in the same direction as the fan  312 , and opposite to the direction that the first motor  328  turns. 
     It will be understood that the propulsion system  305  now comprises a fan  312 , fan shaft  314 , output shaft  318 , ring gear  326 , planet gears  322 , and, additionally, second motor  340 , which rotate in a first direction. The system also includes a sun gear  320 , motor shaft  330  and first motor  328  which spins in a second direction. 
     Consequently, the system  305  can be arranged to cancel the gyroscopic forces as described above. The provision of a second electric motor  340  which rotates in an opposite direction to the first motor  328  allows the designer additional design freedom when choosing motor  328 ,  340  sizes and speeds, which may enable a less compromised design, whilst also substantially cancelling gyroscopic forces. 
       FIG.  4    shows a further alternative propulsion system  405 . The system  405  is similar to the systems  5 ,  305  and comprises a propulsive fan  412  driven by a fan shaft  414 , which is in turn driven by a star reduction gearbox  416 , which is in turn driven by a first electric motor  428  via an input shaft  430 . The propulsion system  405  is mounted by a static mounting system  436 ,  438 . Each of the above components are similar to those of the system  5 , and so are not described in further detail. 
     The propulsion system  405  further comprises second and third motors  442 ,  444 , which are each coupled to the gearbox  416  via respective planet gears  422 , and so rotates in the same direction as the fan  412 , and opposite to the direction that the first motor  428  turns. 
     It will be understood that the propulsion system  405  now comprises a fan  412 , fan shaft  414 , output shaft  418 , ring gear  426 , planet gears  422 , and, additionally, second and third motors  442 ,  444 , which rotate in a first direction. The system also includes a sun gear  420 , motor shaft  430  and first motor  428  which spins in a second direction. It will be appreciated that additional motors could be provided coupled to each planet gear, and configured to rotate with each planet gear, where additional planet gears are present. 
     Again, the system  405  can be arranged to cancel the gyroscopic forces as described above. The provision of second and third electric motors  442 ,  444  which rotate in an opposite direction the first motor  428  again allows the designer additional design freedom when choosing motor  428  sizes and speeds, which may enable a less compromised design, whilst also substantially cancelling gyroscopic forces. 
     Accordingly, a propulsion system is provided having reduced structural loads during manoeuvres and breakages, thereby allowing for reduced structural weight. A designer may design such a propulsion system according to the following design methodology, as set out below. 
     The designer may design one of a propulsor, a motor and a reduction gearbox coupled to the motor at an input side, and the propulsor at an output side. As noted above, the reduction gearbox is configured to provide a reduction ratio between the input and output, and is configured such that the input side rotates in an opposite direction to the output side (e.g. a star gearbox). 
     The designer then modifies the design of the motor, propulsor and reduction gearbox such that gyroscopic forces of the propulsor, motor and reduction gearbox generated during aircraft manoeuvres and/or propulsion system failures are substantially cancelled. This is ensured by changing the diameters, masses and mass distributions of the rotating components, and the gear ratio of the reduction gearbox, until the desired gyroscopic force cancellation is achieved. 
     The disclosed arrangement may also be applied to a parallel propulsion system  505 , such as that shown in  FIG.  8   . 
     The propulsion system  505  of  FIG.  8    comprises a propulsor  512 , which is coupled to a fan shaft  514 , star gearbox  516 , motor shaft  530  and motor  528 , which are arranged in a similar manner to the first embodiment. Also coupled to the motor shaft  520  is a prime mover comprising an internal combustion engine in the form of a gas turbine engine  560 . The gas turbine engine is a single-spool gas turbine engine  560  comprising a compressor  562 , combustor  564  and turbine  566  which are arranged in the normal manner. The compressor  562  supplies compressed air to the combustor  564 , which heats the air, and provides high temperature air to the turbine  566 . The turbine  566  then drives the compressor  562  and the motor  528  via the motor shaft  530 . 
     Consequently, the fan  512  is driven by one or both of the gas turbine engine turbine  566  and the motor  528 . As will be understood, the propulsion system  505  comprises additional rotating components on the input side compared to the propulsion systems  5 ,  305 ,  405 , namely the compressor  562  and turbine  566 , which each rotate in the same direction as the shaft  530 . 
     In this embodiment, the gearbox reduction ratio, gearbox component, fan, gearbox, motor  28 , compressor and turbine masses and mass distributions are arranged such that the gyroscopic forces imposed on the system by the rotating masses on the input side (i.e. the turbine  566 , compressor  562 , motor  528 , shaft  530  and sun gear) are cancelled by the masses on the output side (planets, ring gear, shaft  514  and fan  512 ). 
       FIG.  9    shows a variation of the parallel hybrid propulsion system of  FIG.  8   . 
     In the propulsion system  605  of  FIG.  9   , a motor  628  is provided on the output side of a gearbox  616 , while the gas turbine engine  660  is provided on the input side. As will be understood, the gas turbine engine rotating components now rotate in an opposite direction to the motor  628 . Additionally, the motor  628  rotates at the same speed as the fan  612 . In view of the additional rotating mass on the output side, and reduced rotating mass on the input side, this may help balance the angular momentum of the input and output sides relative to the arrangement shown in  FIG.  8   , which may assist the designer in configuring the propulsion system  605  to have substantially no gyroscopic forces. 
       FIG.  10    shows a further variation of a parallel hybrid propulsion system. 
     The propulsion system of  FIG.  10    is similar to that of  FIG.  9   , and so only differences are described. 
     In the system of  FIG.  10   , the gas turbine engine  760  comprises a two-spool contra-rotating gas turbine engine. The gas turbine engine comprises a low-pressure spool comprising a low-pressure compressor  770 , a low-pressure shaft  730  and a low-pressure turbine  772 . The low-pressure shaft  730  is coupled to an input side of a star reduction gearbox  716 . A motor  728  and fan  712  are coupled to the output side of the gearbox  716 . Additionally, a high-pressure spool is provided, comprising a high-pressure compressor  774 , a high-pressure shaft  776  and a high-pressure turbine  778 . The two spools contra-rotate, thereby at least partially offsetting the gyroscopic forces generated by the engine  760  during manoeuvres. 
     Again, by arranging the clockwise in use rotating components and the anti-clockwise rotating components, along with the reduction gear ratio, such that the gyroscopic forces are substantially cancelled, manoeuvring forces can be substantially reduced. 
     It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein. For instance, the motor could be provided on the input side of the propulsion system having a two-spool gas turbine engine. Any of the motors of  FIGS.  5  to  7    could be employed with any of the motors of any of the propulsion systems of any of the embodiments. Similarly, additional motors installed in locations as shown in  FIGS.  3  and  4    could be provided in the arrangements shown in  FIGS.  8  and  9   . 
     Different types of motors could be employed, such as induction motors. Different types of prime movers, such as piston engines could be employed in place of the gas turbine engine. The gas turbine engine could be omitted, with the aircraft being wholly battery or fuel cell powered, with the prime mover comprising the motor. 
     Similarly, different types of gearboxes could be employed. For example, the gearbox could comprise a step-aside gearbox, a differential gearbox, or any suitable gearbox in which the input and output rotation direction is reversed. In some examples, the gearbox could comprise a step-up gearbox, in which the output speed is faster than the input speed. In other cases, the input and output speed could be equal. Similarly, the reduction ratio could be greater than the examples given, and several gearbox types could be combined.