COOLING CONCEPT FOR TURBINE BLADES OR VANES

Abstract

A turbine assembly with a hollow aerofoil having a main cavity with an impingement tube, insertable inside the main cavity for impingement cooling of an inner surface of the main cavity, and a platform at a radial end of the hollow aerofoil, and a cooling chamber for cooling the platform arranged relative to the hollow aerofoil on an opposed site of the platform. The cooling chamber is limited at a first radial end by a wall segment of the platform and at an opposed radial second end from a cover plate. The impingement tube extends in span wise direction through the cooling chamber from the platform to the cover plate and restricts a sub-cavity of the main cavity. The wall segment includes an entry aperture for a cooling medium to enter from the cooling chamber of the platform into the sub-cavity of the hollow aerofoil.

Claims

1. A turbine assembly comprising: a basically hollow aerofoil having at least a main cavity with at least an impingement tube, which is insertable inside the main cavity of the hollow aerofoil and is used for impingement cooling of at least an inner surface of the main cavity, at least a platform, which is arranged at a radial end of the hollow aerofoil, and at least a cooling chamber used for cooling of at least the platform and which is arranged relative to the hollow aerofoil on an opposed site of the at least one platform and wherein the at least one cooling chamber is limited at a first radial end by at least one a wall segment of the platform and at an opposed radial second end from at least a cover plate, wherein the impingement tube extends in span wise direction at least completely through the cooling chamber from the platform to the cover plate, wherein the impingement tube restricts a sub-cavity of the main cavity, and wherein the at least one wall segment of the at least one platform comprises at least one entry aperture for a cooling medium to enter through the at least one entry aperture from the at least one cooling chamber of the at least one platform into the sub-cavity of the hollow aerofoil.

2. The turbine assembly according to claim 1, wherein the hollow aerofoil comprises a leading edge and a trailing edge, and wherein the impingement tube is located towards the leading edge of the hollow aerofoil and the sub-cavity of the main cavity is located viewed in direction from the leading edge to the trailing edge downstream of the impingement tube.

3. The turbine assembly according to claim 1, wherein the at least one entry aperture in the at least one wall segment of the at least one platform is covered by an orifice plate for controlling a flow of the cooling medium into the sub-cavity.

4. The turbine assembly according to claim 1, wherein the at least one entry aperture in the at least one wall segment of the at least one platform is an insertion aperture through which the impingement tube extends from the at least one cooling chamber of the at least one platform to the main cavity of the hollow aerofoil.

5. The turbine assembly according to claim 1, wherein the at least one entry aperture in the at least one wall segment of the at least one platform is a separate entry aperture from an insert aperture through which the impingement tube extends from the at least one cooling chamber of the at least one platform to the main cavity of the hollow aerofoil.

6. The turbine assembly according to claim 1, wherein the impingement tube ends at the cover plate in a hermetically sealed manner.

7. The turbine assembly according to claim 1, wherein the impingement tube extends substantially completely through a span of the hollow aerofoil.

8. The turbine assembly according to claim 1, further comprising: at least a further platform, wherein the platform and the at least further platform are arranged at opposed radial ends of the hollow aerofoil and wherein the at least further platform comprises at least a further wall segment that comprises at least one further entry aperture for a cooling medium to enter through the least one further aperture from the at least further cooling chamber of the further platform into the sub-cavity of the hollow aerofoil.

9. The turbine assembly according to claim 1, wherein the impingement tube has at least one communicating aperture to allow a flow communication of cooling medium between the impingement tube and the sub-cavity.

10. The turbine assembly according to claim 1, wherein the hollow aerofoil is a turbine blade or vane.

11. The turbine assembly according to claim 1, wherein the hollow aerofoil comprises a trailing edge- and wherein the trailing edge has exit apertures to allow a merged stream of cooling medium from the at least one cooling chamber, from the impingement tube and from the sub-cavity to exit the hollow aerofoil.

12. The turbine assembly according to claim 1, wherein the at least one cover plate of the at least one cooling chamber of the at least one platform is divided by the impingement tube in at least two sections.

13. The turbine assembly according to claim 1 wherein the assembly is cooled by a first stream of cooling medium which is fed to the impingement tube and by a second stream of cooling medium which is fed first to the at least one cooling chamber and thereafter through the at least one entry aperture to the sub-cavity in series.

14. A gas turbine engine comprising: a plurality of turbine assemblies, wherein at least one of the turbine assemblies is arranged according to claim 1.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0049] The present invention will be described with reference to drawings in which:

[0050] FIG. 1: shows a schematically and sectional view of a gas turbine engine comprising several inventive turbine assemblies,

[0051] FIG. 2: shows a perspective view of a turbine assembly with an impingement tube inserted into an aerofoil of the gas turbine engine of FIG. 1 with an entry aperture in a wall segment of a platform,

[0052] FIG. 3 shows a cross section through a turbine assembly along line III-III in FIG. 2,

[0053] FIG. 4: shows a cross section through the aerofoil along line IV-IV in FIG. 3,

[0054] FIG. 5: shows a cross section through the aerofoil along line V-V in FIG. 3,

[0055] FIG. 6: shows a cross section through a first alternative turbine assembly with a alternatively embodied entry aperture,

[0056] FIG. 7: shows a cross section through the aerofoil along line VII-VII in FIG. 6,

[0057] FIG. 8: shows a cross section through the aerofoil along line VIII-VIII in FIG. 6 and

[0058] FIG. 9: shows a cross section through a second alternative turbine assembly with an alternatively embodied impingement tube.

DETAILED DESCRIPTION OF THE ILLUSTRATED EMBODIMENTS

[0059] In the present description, reference will only be made to a vane, for the sake of simplicity, but it is to be understood that the invention is applicable to both blades and vanes of a gas turbine engine. The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine 64 unless otherwise stated. If used, the terms axial, radial and circumferential are made with reference to a rotational axis 74 of the engine 64.

[0060] FIG. 1 shows an example of a gas turbine engine 64 in a sectional view. The gas turbine engine 64 comprises, in flow series, an inlet 66, a compressor section 68, a combustion section 70 and a turbine section 72, which are generally arranged in flow series and generally in the direction of a longitudinal or rotational axis 74. The gas turbine engine 64 further comprises a shaft 76 which is rotatable about the rotational axis 74 and which extends longitudinally through the gas turbine engine 64. The shaft 76 drivingly connects the turbine section 72 to the compressor section 68.

[0061] In operation of the gas turbine engine 64, air 78, which is taken in through the air inlet 66 is compressed by the compressor section 68 and delivered to the combustion section or burner section 70. The burner section 70 comprises a burner plenum 80 one or more combustion chambers 82 defined by a double wall can 84 and at least one burner 86 fixed to each combustion chamber 82. The combustion chambers 82 and the burners 86 are located inside the burner plenum 80. The compressed air passing through the compressor section 68 enters a diffuser 88 and is discharged from the diffuser 88 into the burner plenum 80 from where a portion of the air enters the burner 86 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 90 or working gas from the combustion is channelled via a transition duct 92 to the turbine section 72.

[0062] The turbine section 72 comprises a number of blade carrying discs 94 or turbine wheels attached to the shaft 76. In the present example, the turbine section 72 comprises two discs 94 each carry an annular array of turbine assemblies 10, which each comprises a basically hollow aerofoil 12 embodied as a turbine blade. However, the number of blade carrying discs 94 could be different, i.e. only one disc 94 or more than two discs 94. In addition, turbine cascades 96 are disposed between the turbine blades. Each turbine cascade 96 carries an annular array of turbine assemblies 10, which each comprises a basically hollow aerofoil 12 in the form of guiding vanes, which are fixed to a stator 98 of the gas turbine engine 64. Between the exit of the combustion chamber 82 and the leading turbine blades inlet guiding vanes or nozzle guide vanes 100 are provided.

[0063] The combustion gas 90 from the combustion chamber 82 enters the turbine section 62 and drives the turbine blades which in turn rotate the shaft 76. The guiding vanes 100 serve to optimise the angle of the combustion or working gas 90 on to the turbine blades. The compressor section 68 comprises an axial series of guide vane stages 102 and rotor blade stages 104 with turbine assemblies 10 comprising aerofoils 12 or turbine blades or vanes 100, respectively. In circumferential direction 106 around the turbine assemblies 10 the turbine engine 64 comprises a stationary casing 108.

[0064] FIG. 2 shows in a perspective view a turbine assembly 10 of the gas turbine engine 64. The turbine assembly 10 comprises a basically hallow aerofoil 12, embodied as a nozzle guide vane 100, with two cooling regions, specifically, an impingement cooling region 110 and a fin-pin/pedestal cooling region 112. The former is located at a leading edge 42 and the latter at a trailing edge 44 of the aerofoil 12. At two radial ends 22, 22′ of the hollow aerofoil 12, which are arranged opposed towards each other at the aerofoil 12, a platform and a further platform, referred to in the following text as an outer platform 20 and an inner platform 20′, are arranged. The radial location is defined with the radial direction which in turn is defined in respect to an axis of rotation of the shaft 76 arranged in a known way in the gas turbine engine 64. The outer and the inner platform 20, 20′ both comprise a wall segment 28, 28′, which are oriented basically perpendicular to a span wise direction 34 of the aerofoil 12. Each wall segment 28, 28′ has an insertion aperture 48, which provides access to the aerofoil 12 (only the insertion aperture of wall segment 28 could be seen in FIG. 2). In a circumferential direction 106 of a not shown turbine wheel several aerofoils 12 could be arranged, wherein all aerofoils 12 where connected through the inner and the outer platforms 20, 20′ with one another.

[0065] As could be seen in FIG. 3 that shows a cross section of the turbine assembly 10 along line III-III in FIG. 2, the outer platform 20 and the inner platform 20′ each comprises at least one cooling chamber 24, 24′ referred in the following text as first cooling chamber 24 and a further second cooling chamber 24′. The first and second cooling chambers 24, 24′ are used for cooling of the outer and the inner platforms 20, 20′ and are arranged relative to the hollow aerofoil 12 on opposed sites of the outer and the inner platforms 20, 20′ or their wall segments 28, 28′. The wall segment 28, 28′ of the platform 20, 20′ is a wall separating the cooling chamber 24, 24′ of the platform 20, 20′ from the main cavity 14 of the aerofoil 12 (see below). Thus the wall segment 28, 28′ restricts the main cavity 14 in radial direction. It extends basically perpendicular, advantageously perpendicular, to the span wise direction 34 of the aerofoil 12.

[0066] Both cooling chambers 24, 24′ are limited at a first radial end 26, 26′ by the wall segment 28, 28′ of the outer or the inner platform 20, 20′ and at an opposed radial second end 30, 30′ by a cover plate, referred in the following text as first cover plate 32 and a further second cover plate 32′. The first and second cover plates 32, 32′ are embodied as impingement plates and have impingement holes 116 to provide access for a cooling medium 40 into the first and second cooling chambers 24, 24′.

[0067] A casing 114 of the aerofoil 12 comprises or forms a main cavity 14 spanning the aerofoil 12 in span wise direction 34, wherein the cavity 14 is located in the region of the leading edge 42 or the impingement cooling region 110, respectively. Arranged inside the main cavity 14 is an impingement tube 16, which is inserted via the insertion aperture 48 inside the main cavity 14 during assembly of the turbine assembly 10 for cooling purpose. The impingement tube 16 is used for impingement cooling of an inner surface 18 of the main cavity 14, wherein the inner surface 18 faces an outer surface 118 of the impingement tube 16. The impingement tube 16 extends in span wise direction 34 completely through the cooling chamber 24 from the cover plate 32 to the first platform 20 and it extends in span wise direction 18 along a whole span 50 of the main cavity 14 of the aerofoil 12.

[0068] Moreover, the impingement tube 16 ends at the first cover plate 32 in a hermetically sealed manner, thus preventing a leakage of cooling medium 40 from the impingement tube 16 into the first cooling chamber 24. At the opposed radial end the impingement tube 16 ends or terminates at the further wall segment 28′ of the inner platform 20′ (nor specifically shown) or is sealed via a sealing means, like a lid, in respect to the second cooling chamber 24′. Thus, an entry of cooling medium 40 from the cooling chamber 24′ of the inner platform 20′ into the impingement tube 16 is prevented.

[0069] The inserted impingement tube 16 is located towards or more precisely at the leading edge 42 or is inserted in such a way inside the main cavity 14 to restrict a sub-cavity 36 of the main cavity 14. The sub-cavity 36 is located viewed in axial direction 120—from the leading edge 42 to the trailing edge 44—downstream of the impingement tube 16 or more towards the trailing edge 44 than the impingement tube 16.

[0070] Furthermore, the wall segments 28, 28′ of the outer and the inner platform 20, 20′ each comprises an entry aperture 38, 38′ for the cooling medium 40 to enter through the entry aperture 38, 38′ from the cooling chambers 24, 24′ of the platforms 20, 20′ into the sub-cavity 36 of the hollow aerofoil 12. The entry apertures 38, 38′ in the wall segments 28, 28′ is a section or clearance of the insertion aperture 48 through which the impingement tube 16 is inserted during assembly or through which it extends from the cooling chambers 24 to the main cavity 14. To control the flow of the cooling medium 40 into the sub-cavity 36 the entry apertures 38, 38′ in the wall segments 28, 28′ are covered by an orifice plate 46 with an orifice 122, which can be seen in FIG. 4 that shows a cross section through the aerofoil 12 along line IV-IV in FIG. 3. A cross section through the aerofoil 14 along line V-V in FIG. 3 is shown in FIG. 5.

[0071] Moreover, to allow the cooling medium 40 traveling the impingement tube 16 to exit the impingement tube 16 it has communicating apertures 52 to allow a flow communication of cooling medium 40 between the impingement tube 16, and the sub-cavity 36.

[0072] During an operation of the turbine assembly 10 the impingement tube 16 provides a flow path 124 for the cooling medium 40, for example air. A compressor discharge flow is fed as a first stream 60 of cooling medium 40 from the compressor section 68 to the impingement tube 16 and as a second stream 62 via the impingement holes 116 of the first and second cover plate 32, 32′ into the first and second cooling chambers 24, 24′. The second stream 62 of cooling medium 40 from the first and second cooling chambers 24, 24′ is then discharged into sub-cavity 36 as a platform cooling flow. Thus, the turbine assembly 10 is being cooled by a first stream 60 of cooling medium 40 which is fed to the impingement tube 16 and by a second stream 62 of cooling medium 40 which is fed first to the first and second cooling chambers 24, 24′ and thereafter to the sub-cavity 36 in series.

[0073] For ejection of the cooling medium 40 from the impingement tube 16 to cool the inner surface 18 of the main cavity 14 it comprise not specifically shown impingement holes. The ejected streams of cooling medium 40 from the cooling chambers 24, 24′ and from the impingement tube 16 merge in a space between the outer surface 118 of the impingement tube 16 and the inner surface 18 of the main cavity 14 as well as in the sub-cavity 36. This merged stream flows to the pin-fin/pedestal cooling region 112 located at the trailing edge 44 and exits the hollow aerofoil 12 through exit apertures 54 in the trailing edge 44 (see also FIG. 2).

[0074] It may be possible to divide the cover plate 32 of the cooling chamber 24 of the platform 20 by the impingement tube 16 in at least two sections 56, 58 to choose selected properties to influence flow patterns of the flow of cooling medium 40.

[0075] In FIGS. 6 to 9 alternative embodiments of the turbine assembly 10 and the impingement tube 16 are shown. Components, features and functions that remain identical are in principle substantially denoted by the same reference characters. To distinguish between the embodiments, however, the letters “a” and “b” has been added to the different reference characters of the embodiment in FIGS. 6 to 9. The following description is confined substantially to the differences from the embodiment in FIGS. 1 to 5, wherein with regard to components, features and functions that remain identical reference may be made to the description of the embodiment in FIGS. 1 to 5.

[0076] In FIG. 6 a cross section through an alternatively embodied turbine assembly 10a is shown. The embodiment from FIG. 6 differs in regard to the embodiment according to FIGS. 1 to 5 in that FIG. 6 shows a turbine assembly 10a with separately embodied entry apertures 38a, 38a′. The entry apertures 38a, 38a′ in wall segments 28, 28′ of inner and outer platforms 20, 20′ are separate entry apertures 38a, 38a′ from an insert aperture 48 through which the impingement tube 16 is inserted or through which the impingement tube 16 extends in the assembled state from a cooling chamber 24 of the platform 20 to the main cavity 14 of the hollow aerofoil 12. The arrangement of the separate entry aperture 30 is shown in FIG. 7 that shows a cross section through the aerofoil along line VII-VII in FIG. 6. A cross section through the aerofoil 14 along line VIII-VIII in FIG. 6 is shown in FIG. 8.

[0077] In FIG. 9 a cross section through a turbine assembly 10b analogously formed as in FIGS. 1 to 5 with an alternatively embodied impingement tube 16b is shown. The embodiment from FIG. 9 differs in regard to the embodiment according to FIGS. 1 to 5 in that the impingement tube 16b extends in span wise direction 34 completely through a first cooling chamber 24 from a first or an outer platform 20 to a first cover plate 32 and completely through a second cooling chamber 24′ from a second or inner platform 20′ to a second cover plate 32′. Furthermore, the impingement tube 16b ends at both its radial or longitudinal ends at the first and second cover plate 32, 32′ in a hermetically sealed manner.

[0078] It would be also possible that the impingement tube extends in span wise direction completely through a second cooling chamber from a second platform to a second cover plate. Thus, the impingement tube ends at its second radial or longitudinal end at the second cover plate in a hermetically sealed manner. The impingement tube extends through the inner platform and terminates at its first radial or longitudinal end at the outer platform. A first radial or longitudinal end of the impingement tube is sealed at the wall segment of the outer platform or via a sealing means in respect to the first cooling chamber (not shown).

[0079] In general it would be also possible to provide only one of the wall segments of the inner or outer platform with an entry aperture to allow the flow communication of the cooling medium from the cooling chambers in the sub-cavity. Hence, cooling medium entering one of the cooling chambers of one of the platforms is not fed to the sub-cavity. To provide an outlet for the cooling medium to exit the respective cooling chamber it may be provided with an exit aperture to feed the cooling medium directly into the gas path at an edge of the respective platform (not shown).

[0080] Further it would be also feasible to provide a first stream of cooling medium to the impingement tube from a first platform and to feed the second stream of cooling medium via the cooling chamber to the sub-cavity from the other platform (not shown). To provide an outlet for the cooling medium to exit the cooling chamber without the flow communication (entry aperture) with the sub-cavity it may be provided with an exit aperture to feed the cooling medium directly into the gas path at an edge of the respective platform (not shown).

[0081] Although the invention is illustrated and described in detail by the preferred embodiments, the invention is not limited by the examples disclosed, and other variations can be derived therefrom by a person skilled in the art without departing from the scope of the invention.