COOLED COMPONENT FOR A GAS TURBINE

Abstract

A cooled component for a gas turbine is disclosed, which by an outer side of a wall delimits hot gas passage of the gas turbine and on an inner side has a device for impingement cooling. The impingement cooling device can include a multiplicity of impingement cooling chambers which are arranged next to each other, operate in parallel, are covered by impingement cooling plates which are equipped with impingement cooling holes, and are impinged upon by cooling air during operation.

Description

RELATED APPLICATION

This application claims priority as a continuation application under 35 U.S.C. §120 to PCT/EP2010/053691, which was filed as an International Application on Mar. 22, 2010 designating the U.S., and which claims priority to Swiss Application 00503/09 filed in Switzerland on Mar. 30, 3009. The entire contents of these applications are hereby incorporated by reference in their entireties.

FIELD

The present disclosure relates to the field of gas turbines, such as a cooled component for a gas turbine.

BACKGROUND INFORMATION

Rotor blades, stator blades, heat shields or other components which are exposed to the hot gas flow of a gas turbine can be intensively cooled in order to cope with the thermal and mechanical loads which occur in the machine during operation. One possibility of cooling is film cooling, in which a cooling medium, such as compressed air from the compressor section of the gas turbine, discharges through openings in the wall of the loaded component into the hot gas passage and forms a cooling film on the surface of the component facing the hot gas.

Another possibility of cooling, which can be used alternatively to or additionally to film cooling, is impingement cooling, in which a pressurized cooling medium flows through an impingement cooling plate which is provided with distributed openings, and the jets which emanate therefrom impinge upon the inner wall—arranged at distance—of the loaded component, and wherein the cooling medium intensively absorbs and dissipates heat from the wall during this impingement.

The effectiveness of the impingement cooling in this case can depend upon the type and distribution of the openings provided for it in the impingement cooling plate, upon the distance of the impingement cooling plate from the wall which is to be cooled, upon the selected continuous flow technique of the cooling medium after the impingement, that is to say the continuous flow of the cooling medium after the cooling operation has been carried out, and generally upon the difference of the pressures of the cooling medium which prevail upstream and downstream of the impingement cooling plate. If the cooling medium, however, flows from the chamber, which is formed between impingement cooling plate and wall which is to be cooled, through passages in the wall into the hot gas flow, the pressure difference is determinately influenced by the static pressure which prevails in the hot gas passage at this point.

The last-named configuration is gathered from FIG. 1. FIG. 1 shows in section a gas turbine component 20 with a body 21, which is thermally highly loaded on the outer side (the lower side in FIG. 1). Formed on the inner side of the body 21 is an impingement cooling chamber 23 which is closed off by an impingement cooling plate 22 which is at a distance from the inner wall. The impingement cooling plate 22 has a multiplicity of impingement cooling holes 25 in a distributed arrangement, through which a cooling medium enters the impingement cooling chamber 23 in a jet-like manner and impinges upon the inner wall of the body 21 (arrows in FIG. 1). After the cooling medium has absorbed heat from the wall while impinging upon it, it flows via one (or more) cooling hole(s) outwards into the hot gas passage.

A cooled stator blade segment of a gas turbine, which is cooled by means of a closed steam cooling circuit, is known from printed publication EP-A2-0 698 723. In this case, steam is fed through an inlet, cools the inner sides of the stator blade segment by impingement cooling and is discharged to the outside again through a separate outlet. The segment is divided into individual chambers by ribs in which impingement cooling plates are arranged in each case. The impingement cooling chambers, however, have no connection to the hot gas passage.

A cooled gas turbine blade, in which the inner platform is cooled by impingement cooling, is known from printed publication WO-A1-02/50402. The impingement cooling region is divided into two zones by a rib in order to reduce the crossflows which impair the cooling and to lower local heat transfer coefficients. Arranged in a region which is common to both zones are film cooling holes through which the cooling air discharges from both zones into the hot gas passage.

With these known solutions, the impingement cooling is not optimally adapted directly to the respective static pressure conditions in the hot gas passage and to the local thermal loading in equal measure.

SUMMARY

A cooled component for a gas turbine is disclosed, which comprises: a wall having an outer side which delimits hot gas passage of a gas turbine, and an inner side; a device for impingement cooling on the inner side, wherein the impingement cooling device has a multiplicity of impingement cooling chambers which are arranged next to each other for operation in parallel and for impringement upon by cooling air; and impingement cooling plates which cover the impingement cooling chambers and which are equipped with impingement cooling holes, wherein the impingement cooling chambers are placed to communicate with hot gas passage via separate film cooling holes, and wherein at least one of density and/or distribution and/or diameter or flow cross section of the impingement cooling holes in the impingement cooling plates of the individual impingement cooling chambers is selected with respect to thermal loading and/or static pressure which will occur during operation on the outer side of the wall.

A gas turbine is disclosed comprising: a cooled gas turbine blade component having a wall with an outer side which delimits hot gas passage of a gas turbine, and an inner side; a device for impingement cooling on the inner side, wherein the impingement cooling device has a multiplicity of impingement cooling chambers which are arranged next to each other for operation in parallel and for impringement upon by cooling air; and impingement cooling plates which cover the impingement cooling chambers and which are equipped with impingement cooling holes, wherein the impingement cooling chambers are placed to communicate with hot gas passage via separate film cooling holes, and wherein at least one of density and/or distribution and/or diameter or flow cross section of the impingement cooling holes in the impingement cooling plates of the individual impingement cooling chambers is selected with respect to thermal loading and/or static pressure which will occur during operation on the outer side of the wall.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be described in more detail based on exemplary embodiments in conjunction with the drawings. All elements which are not essential for the direct understanding of the embodiments have been omitted. Like elements are provided with the same designations in the different figures. In the drawings:

FIG. 1 shows a section of a known cooled gas turbine component with a through-chamber for impingement cooling which is covered by a conformable impingement cooling plate;

FIG. 2 shows in perspective view a blade platform with a plurality of separate impingement cooling chambers operating in parallel, according to an exemplary embodiment; and

FIG. 3 shows in a simplified view a section in plane III-III through the blade provided with impingement cooling plates according to FIG. 2.

DETAILED DESCRIPTION

A cooled component for a gas turbine is disclosed wherein a uniform cooling—which is optimally adapted to the local thermal loads and static pressure conditions—of the loaded surface can be achieved.

According to exemplary embodiments, provision is made for a plurality of impingement cooling chambers which operate in parallel and in communication with the hot gas passage by separate film cooling holes, and the density and/or the distribution and/or the diameter of the impingement cooling holes or the flow cross section of the openings in the impingement cooling plates of the individual impingement cooling chambers is, or are, adapted to the respective thermal loading and/or to the respective static pressure on the outer side of the wall which prevails during operation.

An exemplary feature includes a component formed as a blade, such as a stator blade, which is provided with a platform, of the gas turbine, and the impingement-cooled wall is a wall of the platform. Another exemplary feature is a platform integrated into a sequential cooling of the blade, in which the platform and the blade airfoil of the blade are exposed to throughflow by the same cooling medium.

A further feature is a blade that comprises a blade airfoil with a leading edge and a trailing edge, and an impingement cooling chamber, which is arranged downstream of the trailing edge, designed for intensified cooling, by its impingement cooling plate, which is matched to the increased thermal loading as a result of the wake turbulence which develops at the trailing edge. The impingement cooling chamber, which is arranged downstream of the trailing edge, may also be located in a heat shield.

A further feature is a component formed as a heat shield, where the impingement cooling is arranged in different chambers.

Referring to FIGS. 2 and 3, an exemplary embodiment of an impingement-cooled component in the form of a blade 10 of a gas turbine is reproduced. The blade 10 has a blade airfoil 15 extending in the blade longitudinal direction, to which a transversely extending platform 11 is connected at one end. In FIG. 2, the view from the platform 11 from the rear is shown, and in FIG. 3 the section through the platform 11 along the line III-III in FIG. 2 is shown. The platform 11 has a wall 28 which is oriented essentially perpendicularly to the longitudinal direction of the blade and by which the platform adjoins the hot gas passage 27 of the gas turbine. In the hot gas passage 27, the hot gas impinges upon the rotor blades and stator blades which are arranged there, as is indicated in FIG. 3 by the horizontal arrows.

Upstream, the blade airfoil 15 of the blade 10 is delimited by a leading edge 16, and downstream, is delimited by a trailing edge 17. The blade airfoil has a convexly curved suction side and a concavely curved pressure side, which are not directly apparent in the figures. In the case of highly thermally loaded blades, the inner space 12 of the blade airfoil 15 is customarily traversed by a plurality of cooling passages through which a cooling medium, such as cooling air, is directed. On account of the flow conditions in the hot gas passage 27 and on the blade airfoils 15, different pressure and temperature conditions arise around the blade airfoil 15, which are indicated in FIG. 3 by a (static) pressure p_a and a temperature T_a being included on the wall 28 of the platform 11 in the region of the leading edge 16, and in the region of the trailing edge 17 a (static) pressure p_b and a temperature T_b, which as a rule to some extent also differ greatly from the pressure p_a and the temperature T_a, being included.

In proportion to these locally varying thermal conditions, the thermal loading of the wall 28 of the platform 11 on the outer side facing the hot gas passage 27 is also locally different. If the local temperature is high, the thermal loading is can also be high, and vice versa. Furthermore, the local flow conditions play a role because the heat transfer between the hot gas and the wall depends upon whether the local flow happens to be laminar or turbulent, or whether there is even a static zone at the local point.

Exemplary embodiments can therefore provide an ability to correspondingly differently cool thermally differently loaded regions of the wall 28 in order to achieve a cooling or temperature distribution of the platform 11 which is as uniform as possible on the one hand and on the other hand to consume as little cooling air as possible since the consumption of cooling air has an influence upon the efficiency of the machine. For this purpose, the impingement cooling on the rear side of the wall 28 is divided specifically into different regions which have a separate impingement cooling chamber 13 or 13a, 13b in each case, wherein the configuration in respect to cooling technology of these impingement cooling chambers reflects the respective load relationship. Associated with each of the impingement cooling chambers 13a, 13b, which are separated from each other by means of partitions 14 or ribs and operate in parallel, is a separate impingement cooling plate 19a, 19b in which provision is made for impingement cooling openings 18a, 18b of different number and/or distribution and/or of different diameter or flow cross section. The impingement cooling plates 19a, 19b may be individual separate plates, but they may also be designed so that different areas are covered by means of one large common impingement cooling plate so that this impingement cooling plate then stretches across all the impingement cooling chambers 13a, 13b.

The cooling air, which enters the impingement cooling chambers 13a, 13b in jets via the impingement cooling holes 18a, 18b, impinges upon the inner side of the wall 28, absorbs heat from the wall in the process and transports it away by this cooling air then discharging from the impingement cooling chambers 13a, 13b into the adjacent hot gas passage 27. Discharging of the heated cooling air into the hot gas passage 27 is carried out separately for each impingement cooling chamber 13a, 13b via means for cooling representated as separate film cooling holes 26a, 26b. By means of these film cooling holes 26a, 26b, the heated cooling air reaches the outer side of the wall 28 and forms a film there, cooling the outer side of the wall and protecting it against the hot gas in the hot gas passage 27. By means of these film cooling holes 26a, 26b, the respective impingement cooling chamber 13a, 13b at the same time is linked to the static pressure which prevails in the hot gas passage 27 there. In the example of FIG. 3, the static pressure p_a in the hot gas passage 27 is relevant to the impingement cooling chamber 13a, whereas the static pressure p_b in the hot gas passage 27 is of significance to the impingement cooling chamber 13b.

If the local static pressure in the hot gas passage 27 for an impingement cooling chamber is low, the pressure on the other side of the wall 28 can be lowered in order to limit the mass flow of cooling air through this chamber. At the same time, the pressure drop across the associated impingement cooling plate becomes greater and as a result a higher heat transfer coefficient can be achieved during the impingement cooling. If the local static pressure in the hot gas passage 27 for an impingement cooling chamber is high, however, the pressure on the other side of the wall 28 can be designed to be higher in order to prevent penetration of the hot gas into the chamber. Depending upon the prevailing local static pressure and depending upon the thermal loading of the wall 28 in this region, by means of the separate impingement cooling chambers which are linked to the local pressure, the optimum mass flow of cooling air and optimum cooling can be established so that the efficiency of the machine is not unnecessarily reduced.

The platform 11 can, for example, advantageously be integrated into a sequential cooling of the blade 10. The cooling air in this case can be directed through the inner space 12 of the blade airfoil 15 and then flows through the impingement cooling plates 19a, 19b into the impingement cooling chambers 13a, 13b in order to discharge through the film cooling holes 26a, 26b into the hot gas passage 27 after impingement cooling of the wall has been carried out.

The locally adapted and optimized impingement cooling can be advantageous in the case of a blade in which an increased thermal loading occurs as a result of a wake turbulence which develops at the trailing edge 17 of the blade airfoil 15. An impingement cooling chamber 13b which is arranged downstream of the trailing edge 17 is then designed for intensified cooling by its impingement cooling plate 19b (increased hole density in FIG. 3) in order to absorb or compensate the increased thermal loading.

Exemplary embodiments as disclosed herein can include a design for cooling of the platform so that consideration is given to the different pressure or thermal conditions as a result of the flow conditions which prevail beneath this platform, whereby a uniform cooling of the entire platform can be achieved. The different pressure conditions beneath the platform are due at least in part to the fact that the blade contour—such as directly beneath the platform—which is impinged upon is exposed to different pressures, which is why it is intended, even in the case of embodiments disclosed herein, to provide cooling of the platform in sections by means of individual impingement cooling chambers, as is gathered from the figures. This then leads to separated air outflows attributable to the chambers, which is why it can be important that the static pressure prevailing there is taken into account with the cooling. The discharge is of an open design. In the case of the geometric configuration, it can be important that the pressure difference is kept as small as possible, but yet is still of a sufficient level for the hot gases not being able to flash back into the chambers. As a result of the individual cooling in sections, being able to significantly reduce the consumption of cooling air is achieved. Overall, exemplary embodiments as disclosed herein can include an individual design of the impingement cooling chambers in dependence upon the thermal conditions which effectively develop beneath the platform.

With exemplary embodiments as disclosed herein, the exemplary advantages include:

• • The consumption of cooling air can be reduced.
  • Even peak loads can be controlled.
  • Operation can become more reliable if there is no soldered joint, between impingement cooling plate and platform.
  • The cooling can be adapted individually to special thermal loads in specific regions of the blade.
  • Above the impingement cooling plate, provision can be made for additional openings for the bypassing of a bow wave on the blade.
  • The cooling can be combined with sequential cooling arrangements, in which an outer platform is first cooled, and then the blade airfoil and an inner platform are cooled in parallel.

It will be appreciated by those skilled in the art that the present invention can be embodied in other specific forms without departing from the spirit or essential characteristics thereof. The presently disclosed embodiments are therefore considered in all respects to be illustrative and not restricted. The scope of the invention is indicated by the appended claims rather than the foregoing description and all changes that come within the meaning and range and equivalence thereof are intended to be embraced therein.

LIST OF DESIGNATIONS

• 10 Blade (gas turbine)
• 11 Platform
• 12 Inner space (blade airfoil)
• 13 Impingement cooling chamber
• 13a,b Impingement cooling chamber
• 14 Partition
• 15 Blade airfoil
• 16 Leading edge
• 17 Trailing edge
• 18a,b Impingement cooling hole
• 19a,b Impingement cooling plate
• 20 Gas turbine component (blade platform, heat shield)
• 21 Body
• 22 Impingement cooling plate
• 23 Impingement cooling chamber
• 24 Film cooling hole
• 25 Impingement cooling hole
• 26a,b Film cooling hole
• 27 Hot gas passage
• 28 Wall

Claims

1. A cooled component for a gas turbine, which comprises:

a wall having an outer side which delimits hot gas passage of a gas turbine, and an inner side;

a device for impingement cooling on the inner side, wherein the impingement cooling device has a multiplicity of impingement cooling chambers which are arranged next to each other for operation in parallel and for impringement upon by cooling air; and

impingement cooling plates which cover the impingement cooling chambers and which are equipped with impingement cooling holes, wherein the impingement cooling chambers are placed to communicate with hot gas passage via separate film cooling holes, and wherein at least one of density and/or distribution and/or diameter or flow cross section of the impingement cooling holes in the impingement cooling plates of the individual impingement cooling chambers is selected with respect to thermal loading and/or static pressure which will occur during operation on the outer side of the wall.

2. The cooled component as claimed in claim 1, in combination with a gas turbine, wherein the component is a blade provided with a platform, of the gas turbine, and the impingement-cooled wall is a wall of the platform.

3. The cooled component as claimed in claim 2, wherein the platform is integrated into a sequential cooling of the blade, the platform and a blade airfoil of the blade being placed for exposure to throughflow by a cooling medium during operation.

4. The cooled component as claimed in claim 1, wherein the blade comprises:

a blade airfoil with a leading edge and a trailing edge; and

at least one of the impingement cooling chambers, which is arranged downstream of the trailing edge, and designed for intensified cooling by its impingement cooling plate, which is matched to an increased thermal loading due to wake turbulence which develops at the trailing edge.

5. The cooled component as claimed in claim 1, wherein the component is a heat shield.

6. The cooled component as claimed in claim 5, wherein the heat shield is provided with impingement cooling in different impingement cooling chambers.

7. The cooled component as claimed in claim 1, wherein the impingement cooling chambers are directly adjacent one another.

8. The cooled component as claimed in claim 2, wherein the component is a stator blade.

9. A gas turbine comprising:

a cooled gas turbine blade component having a wall with an outer side which delimits hot gas passage of a gas turbine, and an inner side; a device for impingement cooling on the inner side, wherein the impingement cooling device has a multiplicity of impingement cooling chambers which are arranged next to each other for operation in parallel and for impringement upon by cooling air; and impingement cooling plates which cover the impingement cooling chambers and which are equipped with impingement cooling holes, wherein the impingement cooling chambers are placed to communicate with hot gas passage via separate film cooling holes, and wherein at least one of density and/or distribution and/or diameter or flow cross section of the impingement cooling holes in the impingement cooling plates of the individual impingement cooling chambers is selected with respect to thermal loading and/or static pressure which will occur during operation on the outer side of the wall.