GEARED TURBOFAN LOW-PRESSURE TURBINE WITH FLAT HUB

A geared turbofan engine includes a plurality of turbine stages, wherein for each stage (i) of the turbine, an inner radius Ri has a maximum deviation between +1.5% and −3% as compared to the average inner radius of the inner blade platforms of the plurality of stages. The engine further includes a fan, the fan coupled to the turbine stages via a gear.

FIELD OF THE INVENTION

The present invention relates to turbine hubs in a turbine engine.

BACKGROUND INFORMATION

A turbine engine may include a plurality of turbine stages, with each stage having a plurality of rotor blades (e.g. a rotor blade ring) secured for rotation about a rotor axis by a corresponding plurality of turbine hubs, as well as guide vanes between each stage.

One known type of turbine engine includes a turbine module with rotor blade rings that are coupled to the fan via a gearbox. Such an engine is also referred to as a geared turbofan engine (GTF engine). In operation, the fan rotates at a lower speed than the low-pressure turbine module, translated by the gearbox. In such an engine, the “low-pressure turbine module” refers to that section of the aircraft gas turbine of the turbofan engine that is downstream of a most upstream turbine module, which is directly downstream of the combustion chamber, in the direction of flow. The low pressure turbine module can drive a middle and/or innermost shaft of the turbofan engine.

For aerodynamic reasons, the hub contour in low-pressure turbine modules of conventional geared turbofan (GTF) engines generally has a design that initially rises and then falls as illustrated in prior artFIG.1. In particular,FIG.1illustrates the hub configuration of a conventional four stage geared turbofan (GTF) engine, with the radii of the four stages illustrated. As illustrated, the radius rises between the first and second stages before falling between the second, third and fourth stages.

SUMMARY OF THE INVENTION

A disadvantage of the prior art design ofFIG.1is that this hub contour may require comparatively long blade necks (see in particular the third and fourth stages ofFIG.1), which result in increased weight for the engine.

In accordance with an embodiment of the present invention, a geared turbofan engine having a turbine module, in particular a low-pressure turbine (LPT) module for an aircraft engine, is provided which includes a geared turbofan and a flat hub contour. It is to be understood that by geared turbofan it is meant that the fan and the turbine are coupled via a gear. Further, by flat hub contour it is meant that for each stage (i) of the turbine module, the inner radius Riof the hub has a maximum deviation of +1.5% (to the top) and −3% (to the bottom) as compared to an average of the inner radii of the inner blade platforms of the plurality of stages (e.g. R1+R2+R3+R4/4 in the case of a four stage LPT module).

In this regard, it is well known that hubs of a turbine themselves may be contoured or flat, and accordingly when we refer to an inner radius Rinner(or Ri) of a hub, we mean the average value of the radius of the hub to account for the possibility that the hub is a contoured hub.

TMTF (turning mid-turbine frame, or more generally the Inter-Turbine Duct) and TEC (turbine exhaust case) are not part of the control volume.

A geared turbofan with a flat hub contour as defined above provides the advantage that the corresponding blades (or blade necks) are more uniform and short, thus reducing weight. In addition, the use of essentially uniformly short blade necks allows a reduction in the weight of blade roots and disks because with shorter blade necks the rim load (centrifugal force) is reduced, and the rim load is one of the drivers of the weight of the disk.

DETAILED DESCRIPTION

As discussed above, geared turbofan engines conventionally employ a contoured hub profile as illustrated inFIG.1.FIG.1shows a four stage low pressure turbine module of a conventional geared turbofan engine. Illustrated inFIG.1are four stages20,30,40,50each having a bade ring (21,31,41,51). Each blade ring includes a plurality of blades mounted on respective hubs. One of the plurality of blades of each blade ring of each stage (22,32,42,52) is illustrated inFIG.1along with its respective hub (23,33,43,53). The blade rings (which are each comprised of a plurality of blades and hubs) rotate about rotor axis1.1. Also illustrated for completeness are the stationary guide vanes24between each stage. An inner radius Ri of the each hub of stage (i) is illustrated for stages 1-4 and identified as R1, R2, R3and R4. In this regard Ri extends from the axis1.1to the base hub surface on which the blade rests. In this respect, Ri is to be defined as the average radius to account for the possibility that the hub is not flat, but angled or contoured relative to the axis1.1. As can be readily seen fromFIG.1, the radius Ri increases and then decreases, such that R1<R2, R2>R3, and R3>R4. Further it can be seen as R decreases, the length of the blades increases.

FIG.2shows an embodiment in accordance with the present invention.FIG.2shows a four stage low pressure turbine module, including four stages200,300,400,500each having a bade ring (210,310,410,510). Each blade ring includes a plurality of blades mounted on respective hubs. One of the plurality of blades of each blade ring of each stage (220,320,420,520) is illustrated inFIG.2along with its respective hub (230,330,430,530). Also illustrated for completeness are the stationary guide vanes240between each stage. The blade rings (which are each comprised of a plurality of blades and hubs) rotate about rotor axis1.1. An inner radius Ri of the each hub of stage (i) is illustrated for stages 1-4 and identified as R1, R2, R3and R4. In this regard Riextends from the axis1.1to the base hub surface on which the blade rests. In this respect, Riis to be defined as the average radius to account for the possibility that the hub is not flat, but angled or contoured relative to the axis1.1. As can be readily seen fromFIG.2, the radius Riremains flat as contrasted with prior artFIG.1, and the length of the blades for R3and R4are shorter inFIG.2as compared to prior artFIG.1. In each stage (i) of the turbine module, the inner radius Riof the hub has a maximum deviation of +1.5% (to the top) and −3% (to the bottom) as compared to an average of the inner radii (R1+R2+R3+R4/4) of the inner blade platforms of the four stages of the LPT module.

Preferably, in each hub, all radii at the base hub surface (for example, the groove base) are identical within a tolerance of +/−5%.

Preferably, the flat hub contour of the present invention is provided in connection with a turbofan engine with a low-pressure turbine module which has no more than 4 stages, and which includes a fan which is coupled to the low-pressure turbine module via a gearbox. Preferably, the geared turbofan engine has a maximum thrust of at least 70 kN and at most 300 kN (or at most 170 kN, 200 kN or 250 kN), and wherein the low-pressure turbine module has a ratio of annulus area exit to annulus area inlet, Aout/Ain, of at least 2.2 and at most 3.3. The geared turbofan engine preferably has a bypass ratio that is at least 12:1 and at most 15:1. 16. The geared turbofan engine may have a low-pressure turbine rotational speed that is at least 6,500 rpm and at most 14,500 rpm. Further, in operation, the fan may rotate with a maximum thrust between 70 kN and 300 kN, and the low-pressure turbine module may rotate at a rotational speed of 6,500 to 14.500 rpm.

FIG.3shows such a turbofan engine1in a schematic axial section, i.e. viewed in a sectional plane containing its longitudinal axis1.1, and as described in co-pending DE 102022133702 A1, which is hereby incorporated by reference in its entirety. The turbofan engine may be implemented in a regional, short-haul, or medium-haul aircraft, for example.

Functionally, turbofan engine1is divided into compressor2, combustion chamber3and turbine4, the latter having a high-pressure turbine module4.1and a low-pressure turbine module4.2. During operation, air sucked in is compressed in the compressor2, burned in the downstream combustion chamber3with added fuel, and the resulting hot gas is then expanded in the turbine4.

The high-pressure and low-pressure turbine modules4.1,4.2are each constructed in multiple stages, i.e. each have a plurality of guide vanes and rotor blade rings in axial succession. The latter are made to rotate about the longitudinal axis2with the expanding hot gas, with this kinetic energy also being used proportionately to drive the compressor2. A shaft10, on which the rotating blade rings of the low-pressure turbine module4.2rotate, is coupled via a transmission11to a fan12of the turbofan engine1, so the fan12is also driven proportionately with the kinetic energy obtained.

A diameter13of the fan12may for example, be at least 204 cm and at most 247 cm, or at most 219 cm. The low-pressure turbine module4.2, which has 3 or 4 stages, may have an axial length15of at least 20 cm and at most 32 cm, in particular at most 22 cm. The area ratio Aout/Ainof the low-pressure turbine module4.2may be at least 2.2 and at most 3.3, or is at least 2.2 and at most 2.4, and overall an expansion ratio greater than 8 may be realized with the low-pressure turbine module4.2.

The 3-stage low-pressure turbine module may have three rotor blade rings and two or three guide vane rings, with the last (most downstream) ring always being a rotor blade ring. The axial length of a corresponding low-pressure turbine module can, for example, be a maximum of 29 cm (with a possible lower limit of at least 20 cm).

The 4-stage low-pressure turbine module can have four rotor blade rings and three or four guide vane rings, the last ring always being a rotor blade ring. Such a low-pressure turbine module can in particular have an axial length of at most 32 cm (with a possible lower limit of at least 25 cm).

The turbofan engine as a whole can be designed in particular for a thrust range of 110 kN to 140 kN, with the low-pressure turbine module speed (redline) then being between 10,000 and 11,500 rpm.

In accordance with further embodiment s of the present invention, the Rinnerat the fourth stage is smaller than the Rinnerof stages 1-3, while still satisfying a maximum deviation of +1.5% (to the top) and −3% (to the bottom) as compared to each other stage. This is shown inFIG.2. This arrangement provides an advantage in that the inlet Mach number at the last stage is lowered, resulting in improved efficiency

It should also be appreciated that the exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration in any way. Rather, the foregoing detailed description provides those skilled in the art with a convenient road map for implementing at least one exemplary embodiment, it being understood that various changes may be made in the function and arrangement of elements described without departing from the scope of protection as is derived from the claims and the combinations of features equivalent thereto.