US20200263553A1

US20200263553A1 - Cooled turbine rotor blade

Info

Publication number: US20200263553A1
Application number: US16/700,335
Authority: US
Inventors: Hamish SWEIDAN; Keith C. Sadler; Roderick M. Townes
Original assignee: Rolls Royce Deutschland Ltd and Co KG; Rolls Royce PLC
Current assignee: Rolls Royce Deutschland Ltd and Co KG; Rolls Royce PLC
Priority date: 2018-12-04
Filing date: 2019-12-02
Publication date: 2020-08-20

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.

Description

This application claims priority to German Patent Application DE10208130897.3 filed Dec. 4, 2018, the entirety of which is incorporated by reference herein.
The invention relates to a cooled turbine rotor blade for a gas turbine engine, in particular an aircraft turbine engine, and to a rotor-stator stage for a gas turbine providing such a rotor blade as described in the claims.
In modern aircraft engines, it is standard practice to cool the turbine rotor blades that are exposed to high temperature. Turbine rotors with cooled turbine rotor blades are known with different designs of cooling air supply channels, wherein it is referred to EP 1 004 748 B1, EP 1 464 792 B1, and EP 2 551 458 A2 for example.
All these solutions have in common that they aim at providing a maximally effective cooling of the turbine blades with a high-pressure cooling air that is supplied to the turbine rotor in order to minimize the temperature of the turbine blades and to ensure a maximally long service life. As the cited examples show, it is typical that the area of the turbine blade intake at the disk rim of the rotor disk of a turbine rotor is designed in such a manner that a slit-like cooling air supply channel remains between a blade root of a turbine blade and a groove present between the disk fingers of the rotor disk, which support the turbine blade. A cooling air flow is guided through the slit-like cooling air supply channel in the axial direction of the turbine blade and its root portion, and enters cooling channels provided in a aerofoil portion of the blade radially outward of the root portion with the cooling channels branching off radially from the slit-like cooling air supply channel into the blade profile of the aerofoil portion of the turbine blade.
An example for a cooling air inlet on the front side, i.e. the upstream side, of the blade root is shown in EP 3 121 373 A1. Here, the cooling air inlet is designed with a geometry that forms a micro-compressor in order to reduce pressure losses of the cooling air at the entrance into the cooling air supply channel.
For efficient cooling of the blade it may also be desirable to have an inlet for cooling in the rear, i.e. on the axial downstream side, of the blade. However, experiments have shown undesired turbulence when the cooling air boards the spinning blade. The turbulence cause pressure loss, which in turn limits the amount of driving pressure available to push the coolant through the blade. This limits the extent of cooling possible, which in turn limits the maximum operating temperature, thereby limiting both the life of the blade and the engine efficiency.
It is an object to form a cooled turbine rotor blade for a gas turbine engine, in particular an aircraft turbine engine, and a turbine rotor of the kind as it has been described above in such a manner that an effective cooling of the turbine blades and of the disk rim and thus high operating temperatures as well as a long service life of the turbine blades and a high engine efficiency can be ensured, when cooling air enters the spinning turbine rotor blade at its rear side.
The object is achieved with a cooled turbine rotor blade according to the features of patent claim 1 and with a rotor-stator stage for a gas turbine according to the features of patent claim 10.
Further features, advantages and measures are listed in the sub-claims. The features and measures listed in the sub-claims can be combined with one another in advantageous ways.
According to an aspect, there is provided 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 (during operation) of the rotor blade and curved into a radially outward direction.
In use, the blade may have an upstream side and a downstream side defined relative to a flow direction through the gas turbine. The upstream side may correspond to the leading edge of the blade and defines the front side of the rotor blade, while the downstream side may correspond to the trailing edge of the blade and defines the rear side of the rotor blade.
The air pressure at the rear is lower than the front due to the pressure drop across the rotor stage. By using the low pressure side as the cooling air source, less compressor work is required to deliver the cooling air, so the overall engine efficiency is improved. Due to reduced compression, the cooling air is also cooler, so the blade operating temperature can be increased and/or less cooling flow can be used, both resulting in further engine efficiency improvements.
The arrangement of the inlet part of the cooling air supply channel on the rear side of the rotor blade being angled into the rotation direction of the rotor blade has the advantageous effect of a smooth cooling air boarding without additional pressure losses, and therefore, results in an efficient turbine blade cooling. Since the rotor blade is in operation spinning faster than the cooling air, the inlet part of the cooling air supply channel may be profiled to cause a rise of pressure, enabling use of more advanced aerofoil cooling techniques. In this way, the service life of the turbine blade and of the turbine rotor can be considerably increased while high operating temperatures can be used.
In an embodiment, the inlet part of the cooling air supply channel may provide a serpentine design that curves smoothly between the inlet and the cooling air channel, and that may form at least substantially an S-shape. For example, the S-shape may have a first bend after boarding the cooling air, and a second bend leading into a cooling channel radially outwards, with each bend having a smooth angle of more than 90°. With such a design, the inlet part of the cooling air supply channel smoothly turns the cooling air into the cooling air supply channel with minimal pressure losses.
Further, the inlet part of the cooling air supply channel may comprise a geometry that forms a compressor, e.g. with widening diffuser-like diameter, in order to prevent pressure losses. In an expedient embodiment, the inlet of the cooling air may provide a channel geometry that expands in flow direction in a diffuser-like manner. In this way, a compressor is created in a constructionally simple manner. The deceleration of the flow velocity that is taking place inside the compressor leads to a steady increase in static pressure, which ultimately leads to a better cooling system.
Having the described low loss boarding of cooling air available opens up an opportunity to use a low pressure coolant source for high temperature turbines, in particular the air cavity between a rotor stage of a gas turbine and a stator stage following downstream.
The platform may provide at the rear side of the rotor blade a rearward projecting overhang portion forming an air flow discourager, wherein the inlet opening of the cooling air supply channel is located radially inwards of the overhang portion.
The inlet opening position may be positioned at the platform radially outwards of the root portion and optionally of an axial securing device for the root portion at the radial inward side of the overhang portion of the blade platform.
The rear flow discourager of a rotor blade typically forms sealing features of a seal between a rotor stage of a gas turbine and a following stator stage which may combine or interact with neighbouring surfaces for example of neighbouring vanes to form a seal when assembled. The air cavity feeding the blade may be separated from the annulus rim by just the discourager, so the pressure is roughly the same as in the annulus.
However, the disclosure is not limited to such a positioning of the inlet opening of the cooling air supply channel. The person skilled in the art will make use of the advantages of the disclosure depending on the application case, also with a higher or lower radial position of the inlet of the cooling air supply channel.
Further, the rotor blade may have a root portion that is shaped with a fir-treelike profile, so that at least one additional axial cooling air channel may be provided in the root portion or in a gap formed between the root portion and a disk finger groove holding the root portion of the rotor blade on a turbine rotor disk. From such axial cooling air channel one or more radial cooling passages may branch into the aerofoil portion of the turbine blade.
The aerofoil portion may include aerodynamic surfaces that, in operation, are gas-washed by the working fluid, i.e. air, such as a suction surface, a pressure surface, and a platform from which the suction surface and pressure surface extend. The rotor blade may optionally include a shroud at its tip. The shroud may include sealing features to prevent or reduce over-tip leakage flow. The shroud may extend around the entire outer circumference of the blades. Alternatively, the blade may have a partial shroud, or winglet, at its tip (for example extending around a portion of the segment between the blades).
To cause fluid to flow into the blade, the outlet opening of the cooling air supply channel is best to be located high up the blade e.g. at the tip of the aerofoil portion so the rotating cooling air supply channel has a centrifugal pumping effect.
For most effective pumping of cooling air a radially straight up extending design for the cooling air channel may be provided. However, due to the minimized pressure loss at the entering of cooling air there is enough pressure to push the cooling air also through a more complicated blade cooling system. Such a blade cooling system may provide an internal multipass passage with pedestals or turbulators, and throughtrailing-edge ejection holes. Each pass may be arranged to carry cooling air in either a substantially radially outward direction or a substantially radially inward direction. The outlet opening position may be after at least one pass through the internal cooling passage from the cooling air inlet opening. For example, the outlet position may be after two passes through the internal cooling passage, e.g. one pass in the radially outward direction and one pass in the radially inward direction. The radially outward and radially inward directions may be aligned with a longitudinal direction, or chordwise direction, of the blade or the aerofoil portion thereof. The radially outward direction may correspond to a root-to-tip direction of the aerofoil. The radially outward and radially inward directions may refer to the directions relative to a gas turbine engine (for example an axial flow gas turbine engine), for example when the aerofoil is installed in the gas turbine engine.
Again, this augments the extent of cooling possible, which in turn increases the maximum operating temperature, thereby increasing both the life of the blade and the engine efficiency.
According to an aspect, there is provided a rotor-stator stage for a gas turbine, wherein the rotor stage comprises at least one rotor blade according to any one of the preceding claims and an air cavity radially inwards of sealing features between the rotor stage and a neighbouring downstream stator stage forms the source for feeding the rotor blade with cooling air.
Such a rotor-stator stage may provide any one or more of the advantages described herein.
The suggested turbine rotor blade as well as the suggested rotor-stator stage are not necessarily limited to aero engines, their benefits can be realised on any gas turbine e.g. in marine or energy applications.
Apart from the mentioned combination of features, the features specified in the patent claims as well as the features specified in the following exemplary embodiment are suitable respectively on their own or in any combination with each other to embody the subject matter according to the disclosure.

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:

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;

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

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

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.

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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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

10 Gas-turbine engine
11 Air intake
12 Fan
13 Intermediate-pressure compressor
14 High pressure compressor
15 Combustion equipment
16 High pressure turbine
17 Intermediate pressure turbine
18 Low pressure turbine
19 Exhaust nozzle
20 Rotor stage
21 Stator stage
22 Bypass duct
23 Bypass exhaust nozzle
25 Turbine rotor blade
26 Cooling air supply channel
30 Root portion
31 Platform
32 Aerofoil portion
33 Tip
34 Turbine disc
35 Attachment profile, fir tree profile
36 Rear side
37 Overhang portion
38 Axial securing device
39 Seal
40 Inlet opening
41 Outlet opening
42 Inlet part
100 Cooling air
A First air flow
B Second air flow
P View arrow
X Axis

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.