Patent Description:
Gas turbine engines, in particular aircraft engines with geared turbofan engines, require a suitable support for shaft arrangements driving the gearbox and / or the propulsive fan. One shafting arrangement of a geared turbofan engine is described in <CIT>.

European patent publication <CIT> discloses a gas turbine engine comprising a low pressure spool and a reduction gear train. The gear train comprises a sun gear, a carrier having a plurality of planet gears attached thereto, and a ring gear. One of the sun gear, carrier and ring gear is connected to the low pressure shaft, and another of the sun gear, carrier and ring gear provides an output drive. Furthermore, a propulsive fan is mounted fore of the gear train, which is driven by a fan shafting arrangement comprising a fan shaft. The fan shaft is connected to the output drive of the gear train and a fan support shaft which passes through the centre of the gear train along the axis of rotation of the gearbox and fan, wherein the fan shafting arrangement is rotatably supported by at least two axially separated bearings.

US patent publication <CIT> discloses a mounting system for a planetary gear train in a gas turbine engine. The gear train comprises a support strut, a deflection flange and a deflection limiter. The support strut extends between a stationary engine case and a rotating engine shaft that provides input to the planetary gear train. The deflection flange extends from a rotating output component of the gear train. The deflection limiter is connected to the support strut and engages the deflection flange when the gear train becomes radially displaced.

US patent publication <CIT> discloses a turbofan engine having a fan shaft coupling a fan drive gear system to the fan. A low spool comprises a low pressure turbine and a low shaft coupling the low pressure turbine to the fan drive gear system. A core spool comprises a high pressure turbine, a compressor and a core shaft coupling the high pressure turbine to the core spool compressor. A first bearing engages the fan shaft, the first bearing being a thrust bearing. A second bearing engages the fan shaft on an opposite side of the fan drive gear system from the first bearing, the second bearing being a roller bearing. A third bearing engages the low spool shaft and the fan shaft.

US patent publication <CIT> discloses a turbofan engine comprising a fan. A fan drive gear system is configured to drive the fan. A low spool comprises a low pressure turbine and a low shaft coupling the low pressure turbine to the fan drive gear system. An intermediate spool comprises an intermediate pressure turbine, a compressor and an intermediate spool shaft coupling the intermediate pressure turbine to the intermediate spool compressor. A combustor is between a core spool compressor and a high pressure turbine. A first main bearing engages a static support and a forward hub of the intermediate spool. A second main bearing engages the low shaft and the forward hub.

This issue is addressed by a gas turbine with the features of claim <NUM>.

The gas turbine comprises a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear. Typically, the gearbox device is driven by a low pressure or intermediate pressure turbine of the gas turbine, i.e. the sun gear is connected to the input shaft device.

The gearbox device reduces the rotational speed from the turbine to the propulsive fan towards the front of the gas turbine engine making the overall engine more efficient. As will be described further below, the gearbox devices can have different designs.

Depending on the design of the gearbox device, the planet carrier or the ring gear is connected to the propulsive fan via an output shaft device of the gearbox device. The output shaft device can comprise several parts and is generally a hollow shaft with a cross-sectional shape adapted to the load case and the available space within the engine. Furthermore, there is an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device.

There are two alternatives for the design of the input shaft device. In the first alternative the input shaft device has a high rigidity. This could e.g. be a <NUM>% higher stiffness in the radial direction than the support structure (e.g. a static front cone structure, a first strut, a static structure) of the gearbox device. Alternatively and / or additionally, this could be a stiffness greater or equal to <NUM>% of the axial stiffness.

In the second alternative the input shaft device comprises a means for decreasing the rigidity, in particular a diaphragm section.

The bearing devices may comprise more than one bearing. As will be described below, the bearing devices can be positioned axially very close to the gearbox device.

In one embodiment, the inter-shaft bearing system is located axially within or in front of a low-pressure compressor or an intermediate compressor.

In a further embodiment the rear carrier bearing device comprises at least one roller bearing and / or the inter-shaft bearing system comprises at least one ball bearing. It is e.g. also possible to use a double roller bearing with two parallel rows. Furthermore, it is possible that the bearing device comprises bearings which are set apart a certain distance. Those bearings can be identical (e.g. all roller bearings) or they can have a different design.

According to the invention, the inter-shaft bearing system and / or the rear carrier bearing device is axially adjacent to the gearbox device on the input side. The axial distance in the axial direction measured from the centreline of the gearbox device is between <NUM>,<NUM> and <NUM> times, in particular <NUM>,<NUM> to <NUM> times the inner radius of a seat element for the bearing device. This means that the e.g. part of the bearing devices closest to the centreline of the gearbox device can be positioned on the input side of the gearbox device.

Towards the front of the engine a fan shaft bearing system is radially located between a fan shaft as part of the output shaft device and a static front cone structure, in particular the fan shaft bearing system is being axially positioned within the width of the propulsive fan. The static front cone structure - as an example for general static structure within the gas turbine - is relative at rest to the output shaft device. The loads of the fan shaft bearing system can be transmitted to the static part. In one embodiment, the fan shaft bearing system has an outer diameter between <NUM>,<NUM> to <NUM>,<NUM> times the diameter of the propulsive fan, in particular between <NUM>,<NUM> and <NUM>,<NUM> times the diameter of the propulsive fan.

In a further embodiment, the planet carrier comprising the seat element which is extending axially to the rear of the gearbox device providing a radial seat for inter-shaft bearing system.

Further to the rear of the engine an input shaft bearing system is radially located between the input shaft device and a static rear structure, the input shaft bearing system in particular comprising at least one roller bearing. As in the bearing devices or system described above, the input shaft bearing system can comprise more than one row of bearings, the rows being identical or different. The rows can be axially distanced. Alternatively, a ball bearing could be used at location of the input shaft bearing system and a roller bearing in the inter-shaft bearing system.

The shape of the output shaft device can be adapted to spatial requirements. For providing sufficiently mechanical properties, embodiments of the output shaft device can comprise at least one axial cross-section with a conical, sigmoidal or logarithmical shape. In one alternative the fan shaft can be directly attached to the carrier.

In a further embodiment, the output shaft device comprises a curvic or spline coupling. The coupling could e.g. the form of a bellow shaft to achieve a decoupling of the bending between the output shaft and the gearbox device.

In one embodiment of the gas turbine, the load path for force and / or torque from the driving turbine to the propulsive fan extends from the driving turbine to the propulsive fan via the input shaft device, the gearbox device and the output shaft device. There is no through shaft extending through the gearbox device towards the front.

In one embodiment, the ring gear is rigidly connected to the static front cone structure, as it is the case in epicyclic gearbox devices.

In a further embodiment, a static structure comprises struts for transferring radial loads, in particular with parts of the structure comprising an essential linear structure. Linear in this respect means that the load bearing structure elements and struts are located within an angle of less than <NUM>° measured from a radial inward starting point such as e.g. the fan shaft bearing device. From this starting point radial loads are transferred outwards.

In a further embodiment the fan shaft is torsional stiff.

It is also possible that e.g. the gearbox device comprises an epicyclic gearbox with the ring gear being fixed relative to the other parts of the gearbox device and the output shaft device being connected to the planet carrier.

Alternatively, the gearbox device comprises a planetary gearbox in star arrangement with the planet carrier fixed relative to the other parts of the gearbox device and the output shaft device being connected to the ring gear.

The gas turbine engine comprises a gearbox device that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox device may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and / or gear.

The gas turbine engine as described and / or claimed herein may have any suitable general architecture.

The gearbox device may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox device may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox device may be arranged to be driven by any one or more shafts, for example the first and / or second shafts in the example above.

In any gas turbine engine as described and / or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s).

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or <NUM>% span position, to a tip at a <NUM>% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM>. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM>, <NUM> (around <NUM> inches), <NUM> (around <NUM> inches), <NUM> (around <NUM> inches) cm or <NUM> (around <NUM> inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than <NUM> rpm, for example less than <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from <NUM> to <NUM> (for example <NUM> to <NUM>) may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from <NUM> to <NUM> may be in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm, for example in the range of from <NUM> rpm to <NUM> rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades <NUM> on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip<NUM>, where dH is the enthalpy rise (for example the <NUM>-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM> (all units in this paragraph being Jkg-<NUM>K-<NUM>/(ms-<NUM>)<NUM>). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan case.

The overall pressure ratio of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise may be greater than (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) any of the following: <NUM> N kg-<NUM> s, <NUM> N kg-<NUM> s, <NUM> N kg-<NUM> s, <NUM> N kg-<NUM> s, <NUM> N kg-<NUM> s, <NUM> N kg-<NUM> s or <NUM> N kg-<NUM> s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, <NUM> kN, or <NUM> kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus <NUM> deg C (ambient pressure <NUM>. 3kPa, temperature <NUM> deg C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> or <NUM>. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and / or aerofoil portion of a fan blade described and / or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and / or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and / or an organic matrix composite, such as carbon fibre. By way of further example, at least a part of the fan blade and / or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and / or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and / or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be referred to as a blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. For example, at least a part of the fan blades may be machined from a block and / or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example <NUM>, <NUM>, <NUM>, or <NUM> fan blades.

As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and / or engine at the midpoint (in terms of time and / or distance) between top of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach <NUM>,<NUM> to <NUM>,<NUM>, for example <NUM>,<NUM> to <NUM>,<NUM>, for example <NUM>,<NUM> to <NUM>,<NUM>, for example <NUM>,<NUM> to <NUM>,<NUM>, for example <NUM>,<NUM> to <NUM>,<NUM>, for example <NUM>,<NUM> to <NUM>,<NUM>, for example on the order of Mach <NUM>, on the order of Mach <NUM> or in the range of from <NUM>,<NUM> to <NUM>,<NUM>. Any single speed within these ranges may be the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach <NUM>,<NUM> or above Mach <NUM>,<NUM>.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM> (around <NUM> ft), for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> (around <NUM> ft) to <NUM>, for example in the range of from <NUM> to <NUM>, for example in the range of from <NUM> to <NUM>, for example on the order of <NUM>. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to: a forward Mach number of <NUM>,<NUM>; a pressure of <NUM> Pa; and a temperature of -<NUM> (<NUM>).

As used anywhere herein, "cruise" or "cruise conditions" may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.

In use, a gas turbine engine described and / or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example <NUM> or <NUM>) gas turbine engine may be mounted in order to provide propulsive thrust.

The engine core <NUM> comprises, in axial flow series, a low-pressure compressor <NUM>, a high-pressure compressor <NUM>, combustion equipment <NUM>, a high-pressure turbine <NUM>, a low-pressure turbine <NUM> and a core exhaust nozzle <NUM>.

In use, the core airflow A is accelerated and compressed by the low-pressure compressor <NUM> and directed into the high-pressure compressor <NUM> where further compression takes place. The compressed air exhausted from the high-pressure compressor <NUM> is directed into the combustion equipment <NUM> where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high-pressure and low-pressure turbines <NUM>, <NUM> before being exhausted through the nozzle <NUM> to provide some propulsive thrust. The high-pressure turbine <NUM> drives the high-pressure compressor <NUM> by a suitable interconnecting shaft <NUM>.

The low-pressure turbine <NUM> (see <FIG>) drives the shaft <NUM>, which is coupled to a sun wheel, or sun gear, <NUM> of the epicyclic gear arrangement <NUM>.

Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan <NUM>) respectively and / or the turbine and compressor stages that are connected together by the interconnecting shaft <NUM> with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan <NUM>).

The epicyclic gearbox device <NUM> is shown by way of example in greater detail in <FIG>. Practical applications of a planetary epicyclic gearbox device <NUM> generally comprise at least three planet gears <NUM>.

The epicyclic gearbox device <NUM> illustrated by way of example in <FIG> and <FIG> is of the planetary type, in that the planet carrier <NUM> is coupled to an output shaft via linkages <NUM>, with the ring gear <NUM> fixed. In another embodiment the carrier and the output shaft can be manufactured as one part. However, any other suitable type of epicyclic gearbox device <NUM> may be used. By way of further example, the epicyclic gearbox device <NUM> may be a star arrangement, in which the planet carrier <NUM> is held fixed, with the ring (or annulus) gear <NUM> allowed to rotate. By way of further alternative example, the gearbox device <NUM> may be a differential gearbox in which the ring gear <NUM> and the planet carrier <NUM> are both allowed to rotate.

Purely by way of example, any suitable arrangement may be used for locating the gearbox device <NUM> in the engine <NUM> and / or for connecting the gearbox device <NUM> to the engine <NUM>. By way of further example, the connections (such as the linkages <NUM>, <NUM> in the <FIG> example) between the gearbox device <NUM> and other parts of the engine <NUM> (such as the input shaft <NUM>, the output shaft and the fixed structure <NUM>) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox device <NUM> and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of <FIG>. For example, where the gearbox device <NUM> has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in <FIG>.

Optionally, the gearbox device may drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster compressor).

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in <FIG> has a split flow nozzle <NUM>, <NUM> meaning that the flow through the bypass duct <NUM> has its own nozzle that is separate to and radially outside the core engine nozzle <NUM>. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct <NUM> and the flow through the core <NUM> are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.

In <FIG> a schematic view of the front section of geared turbofan engine <NUM> is shown. The view axially extends from the propulsive fan <NUM> in the front to the low-pressure compressor <NUM> towards the rear.

The drive train comprises an input shaft device <NUM> (e.g. comprising the shaft <NUM> shown in <FIG>), here driven by the not shown low-pressure turbine <NUM>. The input shaft device <NUM> is connected to the sun gear <NUM> of the epicyclical gearbox device <NUM>. The input shaft device <NUM> is essentially a hollow tube device providing good torsional stiffness properties.

The output of the gearbox device <NUM> takes place via the planet carrier <NUM> which is connected with an output shaft device <NUM> which has a portion acting as a fan shaft <NUM>. That portion is rigidly connected with the propulsive fan <NUM>. In an alternative embodiment, the output shaft <NUM> can be replaced by a direct connection of the fan disk <NUM> to the carrier <NUM>.

Therefore, the input torque is transmitted from the input shaft device <NUM> to the sun gear <NUM> of the gearbox device <NUM>, and to some extent to the ring gear mount. The planet carrier <NUM> transmits the output torque (at a reduced rotational speed) to the output gear device <NUM> and eventually to the propulsive fan <NUM>.

It is possible that the shape of the shaft devices <NUM>, <NUM> can be more complex and comprises more than one piece.

The shafting arrangement of the embodiment shown in <FIG> also comprises four bearing systems e.g. for taking the mechanical loads or for locating the propulsive fan <NUM> and the gearbox device <NUM>.

The first bearing to be described is an inter-shaft bearing system <NUM> being positioned radially between the planet carrier <NUM> and the input shaft device <NUM>. This inter-shaft bearing system <NUM> comprises one roller bearing. In alternative embodiments, more than one roller bearing (e.g. double bearings, two bearings of different design) or other bearing designs can be used. It is also possible that different bearings of the inter-shaft bearing system <NUM> are positioned at different locations.

The inter-shaft bearing system <NUM> is, in this embodiment, axially adjacent to the gearbox device <NUM> on the input side. The axial distance between the inter-shaft carrier bearing device <NUM> to the gearbox device <NUM> can e.g. be between <NUM>,<NUM> and <NUM> times the inner radius of the inner radius of a seat element <NUM> for the inter-shaft bearing system. This could be in the range of <NUM> to <NUM> measured from the axial front side of the inter-shaft bearing device <NUM> to a centreline <NUM> of the gearbox device <NUM>.

The radial inner seat of the inter-shaft bearing system <NUM> is on seat element <NUM> extending axially to the rear of the gearbox device <NUM>.

A rear carrier bearing device <NUM> is positioned on the input side of the gearbox device <NUM>.

The fan axial load is transferred via the fan-shaft bearing system <NUM> (roller bearing), via the gearbox device <NUM> and into the input-shaft bearing <NUM> towards the rear. With this arrangement the support structures of the bearings can be reduced.

On the output side of the gearbox device <NUM>, the output shaft device <NUM> only has one bearing system, a fan shaft bearing system <NUM>. The radial inner seat of that bearing system is on the fan shaft <NUM>, being a part of the output shaft device <NUM>. The radial outer seat of the fan shaft bearing system <NUM> is connected to a static front cone structure <NUM>. In the embodiment shown a roller bearing is used in the fan shaft bearing system <NUM>. In alternative embodiments, more than one roller bearing (e.g. double bearings, two bearings of different design) or other bearing designs can be used. It would be possible to install a ball bearing and transfer the axial load to the fan <NUM> via the static front cone structure <NUM>.

In the embodiment described herein the fan shaft bearing system <NUM> can have an outer diameter between <NUM>,<NUM> to <NUM>,<NUM> times the diameter of the propulsive fan <NUM>. This range can be between <NUM> and <NUM>.

In an alternative embodiment, the fan shaft bearing system <NUM> is directly located underneath the propulsive fan <NUM>.

The output shaft device <NUM> in the embodiment shown in <FIG> comprises essentially a cylindrical section adjacent to the output side of the gearbox device <NUM> and under the propulsive fan <NUM> (i.e. the fan shaft section <NUM>). In-between there is a conical section <NUM> linking the two cylindrical sections. Conical in this context means that the axial cross-section in this part of the output shaft device <NUM> is a straight line inclined radially inwards. In other embodiments this linking section can have different shapes than the conic shape in <FIG>.

In the embodiment shown in <FIG> the static front cone structure <NUM> and the static structure <NUM> form together one cavity around the gearbox device <NUM>.

The ring gear <NUM> is rigidly connected to the static front cone structure <NUM> but alternatively, it can be connected to a different static part within the engine <NUM>.

The load path for force and / or torque from the driving turbine <NUM>, i.e. the low-pressure turbine <NUM> to the propulsive fan <NUM> extends via the input shaft device <NUM>, the through shaft <NUM>, the gearbox device <NUM> and the output shaft device <NUM>. There is no through shaft.

In <FIG> a variation of the embodiment shown in <FIG> is described. Reference can be made to the respective description above.

The embodiment shown in <FIG> comprises furthermore a structure comprising a plurality of struts <NUM>, <NUM>, <NUM>, <NUM> for taking loads from the moving parts relatively close to the rotational axis <NUM> towards the radially outer parts of the gas turbine engine <NUM>.

In the front the static front cone structure <NUM> transmits radial loads via a first strut <NUM> spanning the airflow A into the engine core <NUM> and the second strut <NUM>. The section of the front cone structure <NUM> and the struts <NUM>, <NUM> are aligned in a relatively straight, linear arrangement, i.e. the section of the front cone structure <NUM>, and the struts <NUM>, <NUM> are positioned within an angular field of is less than <NUM>° measured from the base of the front cone structure <NUM>, i.e. the fan shaft bearing system <NUM>. With this linear arrangement, loads can effectively be transferred.

A similar structure is located axially further to the rear. Here, the static structure <NUM>, and struts <NUM>, <NUM> are aligned in an essentially vertical arrangement.

The two structures <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> described above transfer loads essentially in a radial direction. The channel for the core airflow A provides some stabilization in an essential axial direction resulting in a meshlike structure. This mesh-like structure is torsionaly stiff.

In the embodiments shown in <FIG> and <FIG> the input shaft <NUM> is shown schematically as a straight shaft, i.e. essentially a hollow tube. It is possible that in an alternative embodiment the input shaft comprises flexibility means such as grooves or meandering sections to provide a defined flexibility in the shaft, though this is not in accordance with the present invention.

This is shown in <FIG>, the non-claimed embodiment being a variation of the one shown in <FIG>. The input shaft device <NUM> comprises a diaphragm section <NUM>, i.e. a folded section of the input shaft device.

Claim 1:
A gas turbine engine, in particular an aircraft engine, comprising:
a turbine (<NUM>) connected via an input shaft device (<NUM>) to a gearbox device (<NUM>) having a sun gear (<NUM>), a planet carrier (<NUM>) having a plurality of planet gears (<NUM>) attached thereto, and a ring gear (<NUM>),
the sun gear (<NUM>) is connected to the input shaft device (<NUM>),
the planet carrier (<NUM>) or the ring gear (<NUM>) is connected to a propulsive fan (<NUM>) via an output shaft device (<NUM>) of the gearbox device (<NUM>), with
a rear carrier bearing device (<NUM>) radially between the planet carrier (<NUM>) and a static structure (<NUM>) on the input side of the gearbox device (<NUM>),
an inter-shaft bearing system (<NUM>) being positioned radially between the input shaft device (<NUM>) and the planet carrier (<NUM>) of the gearbox device (<NUM>).
the input shaft device (<NUM>) having a high rigidity, characterised in that the planet carrier (<NUM>) comprises a seat element (<NUM>) extending axially to the rear of the gearbox device (<NUM>) providing a radial seat for the inter-shaft bearing device (<NUM>), the inter-shaft bearing system (<NUM>) is axially adjacent to the gearbox device (<NUM>) on the input side with an axial distance measured from the centreline (<NUM>) of the gearbox device (<NUM>) between <NUM>,<NUM> and <NUM> times the inner radius of a seat element (<NUM>) for the inter-shaft bearing system (<NUM>).