COOLED TURBINE ROTOR BLADE

Abstract

A cooled turbine rotor blade for a gas turbine engine includes an airfoil and a profiled root radially inward of the airfoil, a platform between the airfoil and the root, and a cooling air supply channel extending through the platform into an interior of the airfoil and therein up to an outlet opening. An inlet opening of the cooling air supply channel is located at the rear side of the rotor blade, and an inlet part of the cooling air supply channel with the inlet opening is angled into the direction of rotation of the rotor blade and curved into a radially outward direction. Further, a rotor-stator stage for a gas turbine includes a rotor blade as above, and an air cavity radially inwards of sealing features between the rotor stage and a neighboring downstream stator stage to form a source of cooling air for the rotor blade.

Claims

1. A cooled turbine rotor blade for a gas turbine engine, in particular an aircraft turbine engine, comprising an aerofoil portion and a profiled root portion radially inward of the aerofoil portion in installed state, a platform between the aerofoil portion and the root portion, and at least one cooling air supply channel that extends through the platform into the interior of the aerofoil portion and therein up to at least one outlet opening, wherein an inlet opening of the cooling air supply channel is located at the rear side of the rotor blade, and an inlet part of the cooling air supply channel with the inlet opening is angled into the direction of rotation of the rotor blade and curved into a radially outward direction.

2. The cooled turbine rotor blade as claimed in claim 1, wherein the inlet part of the cooling air supply channel provides a serpentine design.

3. The cooled turbine rotor blade as claimed in claim 1, wherein the inlet part of the cooling air supply channel forms at least substantially a S-shape.

4. The cooled turbine rotor blade as claimed in claim 1, wherein the inlet part of the cooling air supply channel comprises a geometry that forms a compressor diffuser.

5. The cooled turbine rotor blade as claimed in claim 1, wherein the inlet opening of the cooling air supply channel is located at the platform at its radially inwards side.

6. The cooled turbine rotor blade as claimed in claim 1, wherein the platform provides at the rear side of the rotor blade a rearward projecting overhang portion, wherein the inlet opening of the cooling air supply channel is located radially inwards of the overhang portion.

7. The cooled turbine rotor blade as claimed in claim 1, wherein at least one outlet opening of the cooling air supply channel is located at a tip of the aerofoil portion.

8. The cooled turbine rotor blade as claimed in claim 1, wherein at least one cooling air supply channel is extending radially straight up in the aerofoil portion.

9. The cooled turbine rotor blade as claimed in claim 1, wherein the cooling air supply channel is configured as a multipass passage system.

10. A rotor-stator stage for a gas turbine, wherein the rotor stage comprises at least one rotor blade according to claim 1, and an air cavity radially inwards of sealing features between the rotor stage and a neighbouring downstream stator stage forms a source for feeding the rotor blade with cooling air.

Description

[0030] Other advantages and embodiment possibilities of the cooled turbine blade according to the disclosure also follow from the patent claims and the exemplary embodiment that will be described in principle below by referring to the drawing, in which:

[0031] FIG. 1 shows a portional side view of a gas turbine engine providing cooled turbine rotor blades and a rotor-stator stage according to the disclosure;

[0032] FIG. 2 shows a partial portion view of a rotor-stator stage with a turbine rotor blade providing an internal cooling air supply channel;

[0033] FIG. 3 shows a simplified perspective view the rotor blade of FIG. 2 seen in the direction of arrow P in FIG. 2; and

[0034] FIG. 4 shows a perspective view of a part of the rotor blade of FIG. 2 and FIG. 3 with an inlet opening of a cooling air supply channel.

[0035] With reference to FIG. 1, a ducted fan gas turbine engine 10, here for use in an aircraft, has a principal and rotational axis X-X. The engine 10 comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 12, a bypass duct 22 and a bypass exhaust nozzle 23.

[0036] The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.

[0037] The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 16, 17, 18 respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.

[0038] As the air passes through the gas turbine engine 10 it is heated to high temperatures. In particular, the first airflow B reaches high temperatures as it passes, through the core of the engine. Typically, particularly high temperatures may be reached at the exit of the combustion equipment 15, and as the air subsequently passes through the high, intermediate and low-pressure turbines 16, 17, 18.

[0039] Gas temperatures in the turbine can be in excess of 2100 K. It is desirable to operate the turbine at the highest possible temperature because generally, for a given gas turbine configuration, increasing the turbine entry gas temperature will produce more specific thrust.

[0040] Therefore, in order to cool turbine rotor blades 25, internal cooling air supply channels 26 may be formed within the blades 25. These internal cooling air supply channels 26 allow cooling air to be passed through the blades to remove heat through convection.

[0041] Typically, cooling air 100 may be bled from the compressor 13, 14, prior to combustion, for example from the HP compressor. Typical cooling air temperatures are between 700 and 900 K.

[0042] In addition to the cooling air 100 that is provided to the turbine rotor blade 25 from the compressor 13, 14 in the arrangement shown in FIG. 2 to FIG. 4, more air is required for sealing purposes between a rotor stage 20 and a for example downstream neighbouring stator stage 21.

[0043] The turbine rotor blade 25 may be any type of turbine blade. For example, the turbine blade 25 may be part of a high pressure turbine 16, an intermediate pressure turbine 17, or a low pressure turbine 18. The turbine blade 25 may be part of any type of gas turbine engine, for example a ducted fan gas turbine (turbofan) engine 10 such as that shown in FIG. 1, a turbojet, a turboprop, a turboshaft, an open rotor engine, or any other gas turbine engine, for example axial flow or radial flow.

[0044] The turbine blade 25 has a root portion 30, a platform 31, an aerofoil portion 32, and a tip 33. Cooling air 100 passes through a cooling air supply channel 26 from an inlet opening 40 through the platform 31 into the interior of the aerofoil portion 32 and therein radially outwards up to an outlet opening 41 at the tip 33 of the rotor blade 25.

[0045] The root portion 30 may allow the rotor blade 25 to be attached to a corresponding component of a gas turbine engine, for example to a turbine disc 34. The term root portion 30 may be used to refer to parts of the rotor blade 25 that are radially inward of the platform 31. The root portion 30 has an attachment profile 35 which may be a fir tree as shown in FIG. 4. The attachment profile 35 allows the rotor blade 25 to be attached to the corresponding turbine disc 34.

[0046] The rotor blade 25 may be said to be shrouded. However, different types of blades may be used, for example also partially shrouded turbine rotor blades.

[0047] The inlet opening 40 of the cooling air supply channel 26 is located at the rear side 36 of the rotor blade 25 at the platform 31. The platform 31 as shown comprises at the rear side 36 of the rotor blade 25 a rearward projecting overhang portion 37 forming an air flow discourager. The rearward overhang portion 37 forms sealing features of a seal 39 between the rotor stage 20 and the following stator stage sealing an air cavity 27 from which the cooling air supply channel 26 is feeded.

[0048] In the shown embodiment, the inlet opening 40 of the cooling air supply channel 26 is positioned at the beginning of the overhang portion radially inwards of the overhang portion 37 and radially outwards of the root portion 30 and its axial securing device 38.

[0049] The inlet opening 40 and a following inlet part 42 of the cooling air supply channel 26 are angled into the direction of rotation (in operation) of the rotor blade, and curved into an radially outward direction. Thus, the inlet part 42 of the cooling air supply channel 26 may provide a serpentine design that may form at least substantially a S-shape with a radially straight upward ending. Also more bends than in S-shape and/or sharper turns, e.g. 180 bends, may be chosen by the person skilled in the art.

[0050] With such design, a smooth entry of cooling air 100 into the inlet opening 40 and a pressure recovery within the inlet part 42 of the cooling air supply channel 26 is achieved, providing the afore-mentioned advantages.

LIST OF REFERENCE SIGNS

[0051] 10 Gas-turbine engine [0052] 11 Air intake [0053] 12 Fan [0054] 13 Intermediate-pressure compressor [0055] 14 High pressure compressor [0056] 15 Combustion equipment [0057] 16 High pressure turbine [0058] 17 Intermediate pressure turbine [0059] 18 Low pressure turbine [0060] 19 Exhaust nozzle [0061] 20 Rotor stage [0062] 21 Stator stage [0063] 22 Bypass duct [0064] 23 Bypass exhaust nozzle [0065] 25 Turbine rotor blade [0066] 26 Cooling air supply channel [0067] 30 Root portion [0068] 31 Platform [0069] 32 Aerofoil portion [0070] 33 Tip [0071] 34 Turbine disc [0072] 35 Attachment profile, fir tree profile [0073] 36 Rear side [0074] 37 Overhang portion [0075] 38 Axial securing device [0076] 39 Seal [0077] 40 Inlet opening [0078] 41 Outlet opening [0079] 42 Inlet part [0080] 100 Cooling air [0081] A First air flow [0082] B Second air flow [0083] P View arrow [0084] X Axis