DE102013109151A1 - Blade assembly for a turbomachine - Google Patents

Abstract

Blade assemblies are created. The blade assembly includes a shank and an airfoil positioned radially outward of the shank. The blade assembly further includes a main cooling circuit formed in the airfoil and the shank, the main cooling circuit having seven passages, each of the seven passages being in fluid communication with an adjacent one of the seven passages. A maximum number of rotations in each of the seven passes is less than or equal to about 0.4.

Description

FIELD OF THE INVENTION

The present disclosure relates generally to turbomachines, and more particularly to blade assemblies in turbomachinery.

BACKGROUND TO THE INVENTION

Gas turbine systems are an example of turbomachinery widely used in the fields of, for example, power generation. A conventional gas turbine system includes a compressor section, a combustor section and a turbine section. During operation of the gas turbine system, various components in the system are exposed to high temperature flows that may cause the components to fail. Because higher temperature flows generally result in increased performance, efficiency, and output of the gas turbine system, the components exposed to the high temperature flows should be cooled to allow the gas turbine system to operate at elevated temperatures.

For each stage of the turbine section or hot gas path section of a gas turbine system, many system requirements should be met to achieve design goals, including overall improved efficiency and airfoil loading. In particular, the blades of the various stages of the turbine section should meet the operating requirements for that particular stage and also meet blade cooling range and wall thickness requirements. Internal cooling requirements should be optimized, requiring a unique inner core profile to meet the stage performance requirements that allow the turbine to operate in a safe, efficient and hassle-free manner.

In particular, a current tendency in the design of blades is to increase the maximum thickness (between the pressure side and the suction side) of the blade airfoil to improve aerodynamic efficiency. Further, reductions in airfoil wall and rib thicknesses (between the various cooling passages defined therein) have been promoted to minimize the overall weight of the blade. However, the combination of these two tendencies has led to large, high aspect ratio cooling passages requiring increased cooling flow therethrough to achieve the required thermal convection coefficient for proper cooling thereof. The present inventors have found that the Mach number of the cooling flow through the cooling passages is reduced due to these tendencies, resulting in increased Coriolis effects in these cooling passages. For each cooling passage this results in uneven cooling with an increase in the heat transfer coefficient on either the pressure side or the suction side of the passage and a reduction in the heat transfer coefficient on the other side. As a result, cooling of these blades may be inefficient and uneven and may result in blade damage.

Accordingly, improved blades are desired in the art. In particular, blades with improved cooling circuits that provide more efficient and well-distributed cooling would be advantageous.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in the description which follows, or may be obvious from the description, or may be learned by practice of the invention.

In one aspect, a blade assembly is disclosed. The bucket assembly includes a shank and an airfoil positioned radially outward of the shank. The bucket assembly further includes a main cooling circuit defined in the airfoil and shaft, the main cooling circuit having seven passages, each of the seven passages being fluidly connected to an adjacent one of the seven passages. A maximum rotation number in each of the seven passes is less than or equal to about 0.4.

Particularly advantageous and preferred embodiments of the blade assembly according to the first aspect of the present invention include one or more of the following:
The number of rotations in each of the seven passes is preferably in a range between about 0.01 and about 0.35.

A first pass of the seven passages may have a maximum cross sectional area in the airfoil that is smaller by a size in a range between about 10% and about 20% than a maximum cross sectional area of another passage of the seven passages in the airfoil.

The bucket assembly may further include an outlet passage defined in the airfoil and a seventh Passage of the seven passages is in fluid communication.

The bucket assembly may further include a front cooling circuit and a rear cooling circuit defined in the airfoil and the shaft, respectively.

The main cooling circuit preferably consists of seven passes.

The bucket assembly may further include a platform positioned radially between the shank and the airfoil.

The airfoil may have outer surfaces defining a pressure side and a suction side extending between a leading edge and a trailing edge.

The blade assembly may further include a dovetail. The dovetail may define a root and the main cooling circuit may extend to the root.

In another aspect, a turbomachine is disclosed. The turbomachine includes an inlet section, an outlet section, a compressor section and a turbine section. The turbine section includes a plurality of blade assemblies. Each of the plurality of blade assemblies includes a stem and an airfoil positioned radially outward of the stem. Each of the plurality of blade assemblies further includes a main cooling circuit defined in the airfoil and shaft, the main cooling circuit having seven passages, each of the seven passages being fluidly connected to an adjacent one of the seven passages. A maximum rotation number in each of the seven passes is less than or equal to about 0.4.

Particularly advantageous and preferred embodiments of the turbomachine according to the second aspect of the present invention include one or more of the following:
The number of rotations in each of the seven passes is preferably in a range between 0.01 and 0.35.

A first pass of the seven passages may have a maximum cross-sectional area in the airfoil that is smaller by a size in a range between about 10% and about 20% than a maximum cross-sectional area of another passage of the seven passages in the airfoil.

The turbomachine may further include an outlet passage defined in the airfoil and in fluid communication with a seventh passage of the seven passages.

The turbomachine may further include a front cooling circuit and a rear cooling circuit defined respectively in the airfoil and the shaft.

The main cooling circuit preferably consists of seven passes.

The turbomachine may further include a platform positioned radially between the shaft and the airfoil.

The airfoil may have outer surfaces defining a pressure side and a suction side extending between a leading edge and a trailing edge.

The turbomachine may further include a dovetail. The dovetail may define a root and the main cooling circuit may extend to the root.

These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A broad and enabling disclosure of the present invention, including the best mode thereof, which is directed to one skilled in the art, is provided in the description which refers to the accompanying drawings, in which:

1 a schematic representation of a gas turbine system;

2 FIG. 4 is a sectional side view of the turbine section of a gas turbine system according to an embodiment of the present disclosure; FIG.

3 a perspective view of a blade assembly according to an embodiment of the present disclosure;

4 a front view illustrating the inner components of a blade assembly according to an embodiment of the present disclosure; and

5 FIG. 10 is a plan view illustrating various internal components of a blade assembly according to an embodiment of the present disclosure. FIG.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided to illustrate the invention, not for the purpose of limiting the invention. In fact, it will be apparent to those skilled in the art that various modifications and changes may be made to the present invention without departing from the scope or spirit of the invention. For example, features that are illustrated or described as part of one embodiment may be used in another embodiment to yield a still further embodiment. Thus, it is intended that the present invention cover such modifications and changes as fall within the scope of the appended claims and their equivalents.

1 shows a schematic representation of a turbomachine, which in the illustrated embodiment, a gas turbine system 10 is. The system 10 can be a compressor section 12 , a combustion chamber section 14 and a turbine section 16 contain. The compressor section 12 and the turbine section 16 can over a wave 18 be coupled with each other. The wave 18 For example, a single shaft or may be formed by a plurality of shaft segments which are coupled together to the shaft 18 to build. An inlet section 19 may be an air flow to the compressor section 12 supply, and exhaust gases can from the turbine section 16 via an outlet section 20 spent and omitted and / or in the system 10 or any other suitable system.

The turbine section 16 can contain several turbine stages. For example, the turbine section 16 in one embodiment, have three stages, as in FIG 2 illustrated. For example, a first stage of the turbine 16 a plurality of circumferentially spaced apart vanes 21 and blades 22 contain. The vanes 21 can along the circumference around the shaft 18 be arranged around and fixed. The blades 22 can along the circumference around the shaft 18 arranged around and with the shaft 18 be coupled. A second stage of the turbine section 16 may be a plurality of circumferentially spaced apart vanes 23 and blades 24 contain. The vanes 23 can along the circumference around the shaft 18 be arranged around and fixed. The blades 24 can along the circumference around the shaft 18 arranged around and with the shaft 18 be coupled. A third stage of the turbine section 16 may be a plurality of circumferentially spaced apart vanes 25 and blades 26 contain. The vanes 25 can along the circumference around the shaft 18 be arranged around and fixed. The blades 26 can along the circumference around the shaft 18 arranged around and with the shaft 18 be coupled. The different stages of the turbine section 16 can in the turbine 16 in the path of the hot gas flow 28 be arranged. It should be understood that the turbine section 16 is not limited to three levels, but rather that any number of stages are within the scope and spirit of the present disclosure.

Each of the blades 22 . 24 . 26 can be a blade arrangement 30 have, as in the 3 to 5 illustrated. In exemplary embodiments, the blade assembly is 30 according to the present disclosure, a blade 22 the first stage, although in alternative embodiments, the blade assembly 30 may be any suitable blade. The blade arrangement 30 can generally be a shaft 32 and an airfoil 34 contain. The blade 34 can be radially outside of the shaft 32 be positioned and radially outward of the shaft 32 extend. In addition, the blade assembly 30 a platform 36 contain. The platform 36 can be radial between the shaft 32 and the blade 34 be positioned and can also either the shaft 32 or the blade 34 or both are surrounded. Furthermore, the blade assembly 30 a swallowtail 38 contain. The swallowtail 38 can be radially inward of the shaft 32 extend and, in some embodiments, a radially inner portion of the shaft 32 be. The swallowtail 38 can be used to couple the blade assembly 30 be configured with a (not shown) impeller.

The blade 34 may have outer surfaces, generally a pressure side 42 and a suction side 44 contain and define that are between a leading edge 46 and a trailing edge 48 extend. Furthermore, the blade assembly may 34 between a tip 52 and a root 54 extend. The summit 52 can z. B. a radially outer portion of the airfoil 34 be. The root 54 may be a radially inner portion of the shaft 32 or the swallowtail 38 be.

A scoop arrangement 30 According to the present disclosure further includes one or several cooling circuits defined therein. A refrigeration cycle may generally be through the airfoil 34 and / or the shaft 32 extend. A cooling fluid (usually air or other suitable gas) may be flowed through the cooling circuits to the blade assembly 30 and in particular the airfoil 34 to cool. For example, a blade assembly 30 a main cooling circuit 60 contain. The main cooling circuit 60 can in the airfoil 34 and the shaft 32 be educated. For example, the main cooling circuit 60 through the shaft 32 through to the root 54 extend, and he can further from the root 54 out through the shaft 32 through and into the blade 34 extend into it.

As illustrated, the main cooling circuit contains 60 several cooling passages. In particular, for example, there is a main cooling circuit 60 According to the present disclosure, seven cooling passages including a first cooling passage 61 , a second cooling passage 62 , a third cooling passage 63 , a fourth cooling passage 64 , a fifth cooling pass 65 , a sixth cooling passage 66 and a seventh cooling passage 67 belong. Each cooling passage may be substantially radial through the blade assembly 30 extend. Each cooling passage may be connected to adjacent cooling passages and thus be in fluid communication. For example, the first cooling passage 61 with the second cooling passage 62 in fluid communication with the third cooling passage 63 can be in fluid communication, etc. By itself, the cooling channels in a row between the front edge 46 and the trailing edge 48 aligned, and thus can form a series substantially in the chordwise direction. The cooling passages may be arranged in order, the first cooling passage 61 relatively closest to the trailing edge 48 and the seventh cooling passage 67 relatively closest to the leading edge 46 is positioned. A cooling fluid may pass through each cooling passage in the main cooling circuit 60 stream. Further, the connected cooling passages may have a substantially serpentine pattern so that the flow direction may change through the cooling passages. For example, the general flow direction in the first, third, fifth, and seventh cooling passages 61 . 63 . 65 and 67 in a substantially radially outward direction to the tip 52 point, and the general flow direction in the second, fourth and sixth cooling passage 62 . 64 and 66 may be in a substantially radially inward direction toward the root 54 point out.

Each cooling passage of the main cooling circuit 60 may at least partially in the airfoil 34 and optionally the shaft 32 be defined. One or more cooling passages may additionally include one or more inlet sections through which a cooling fluid to the main cooling circuit 60 is delivered. For example, the first cooling passage 61 as illustrated, a first inlet section 68 and a second inlet portion 69 included in the shaft 32 are trained and get down to the root 54 extend. The inlet sections 68 . 69 For example, in the shaft 32 as illustrated, around a single first cooling pass 61 to build.

The blade arrangement 30 may additionally contain various other cooling circuits. For example, the blade assembly 30 a front cooling circuit 70 contain. The front cooling circuit 70 can in the airfoil 34 and the shaft 32 be educated and can move from the root 54 extend as illustrated. Furthermore, the front cooling circuit 70 in the chordwise direction substantially adjacent to the leading edge 46 such as between the leading edge 46 and the main cooling circuit 60 to be positioned. The front cooling circuit 70 For example, any suitable number of cooling passages, such as, for example, a first cooling passage in exemplary embodiments 71 and a second cooling passage 72 , contain.

The blade arrangement 30 may additionally or alternatively a rear cooling circuit 80 contain. The rear cooling circuit 80 can in the airfoil 34 and the shaft 32 be educated and can move from the root 54 extend as illustrated. Furthermore, the rear cooling circuit 80 in the chordwise direction substantially adjacent to the trailing edge 48 such as between the trailing edge 48 and the main cooling circuit 60 to be positioned. The rear cooling circuit 80 For example, in exemplary embodiments, a first cooling passage may be any number of cooling passages 81 , contain.

In addition, outlet passages 90 in the airfoil 34 be educated. Each outlet passage 90 can with a passage of a cooling circuit, such as the main cooling circuit 60 , the front cooling circuit 70 or the rear cooling circuit 80 , as illustrated, be connected and thus in fluid communication. The outlet passages 90 may be provided to flow the cooling fluid therethrough from the cooling circuits and the cooling fluid from the blade assembly 30 issue. In particular, the outlet passages 90 in the airfoil 34 be educated and with the seventh passage 67 in flow communication, as illustrated, to a cooling flow from the main cooling circuit 60 issue. In some embodiments, the exhaust passages 90 For example, be aligned such that the output cooling fluid film cooling of the blade assembly 30 achieved.

Now again on the main cooling circuit 60 Reference is made, contains such a cycle 60 various features and properties, which advantageously provide cooling of the blade assembly 30 improve. For example, as explained above, the main cooling circuit 60 seven cooling passages. This represents an increase in the typical number of cooling passages included in heretofore known blade assemblies. By increasing the number of cooling passages, the overall weight of the blade assembly can be increased 30 and the thickness of the walls and ribs of the airfoil 34 be reduced. Further, however, the Mach number of the cooling flow in the cooling passages may also advantageously be increased, resulting in reduced Coriolis effects in the cooling passages, thus providing more efficient and evenly distributed cooling of the blade assembly 30 through the cooling passages. For example, the present inventors have found that cooling in a blade assembly is improved by reducing the number of rotations of the cooling flow in the cooling passages to less than or equal to about 0.4. The rotation number is a dimensionless number that is in inverse proportion to the Coriolis force and can be calculated based on the equation: Ro = ω · b | (W | (ρA)), where Ro = rotational number (dimensionless), ω = rotational velocity, b = cross-sectional length of the passage (other than the cross-sectional width a, see 5 ), W = mass flow rate of the cooling fluid, ρ = cooling fluid density and A = cross-sectional area of the cooling passage (eg, a · b for a square or rectangular cooling passage area). The design of the present main cooling circuit 60 according to the present disclosure thus provides a cooling flow with a maximum rotation number in each of the seven cooling passages 61 . 62 . 63 . 64 . 65 . 66 . 67 less than or equal to about 0.4, for example less than or equal to about 0.35 in some embodiments, for example, between about 0.01 and about 0.4 in some embodiments, such as between about 0.01 and 0 in some embodiments , 35, ready.

The present inventors have also found that additional additional design features of the present main refrigeration cycle 60 a cooling of the blade assembly 30 improve. For example, the present inventors have the cross sectional area of the first passage 61 of the main cooling circuit 60 in the airfoil 34 in relation to the cross-sectional area of the remaining passages 62 . 63 . 64 . 65 . 66 . 67 in the airfoil 34 reduced. The cross-sectional area may be in the in 5 illustrated at any suitable location on the airfoil 34 be calculated. The maximum cross-sectional area in the airfoil 34 the first passage 61 Thus, it may be smaller by a size in a range between about 10% and about 20% than a maximum cross-sectional area of any other passage 62 . 63 . 64 . 65 . 66 . 67 in the airfoil 34 , Such a reduction in the first pass 61 in relation to the remaining passes 62 . 63 . 64 . 65 . 66 . 67 further reduces the Korioris effect in the first pass 61 , which is a place where the Korioris effect used to be relatively more pronounced.

The present design of the main cooling circuit 60 also offers several other benefits. For example, cooling in the various turning regions in which cooling flow is rotated between adjacent cooling channels is enhanced due to the increased Mach numbers of the cooling flow as the cooling flow enters the turning regions. Thus, the present blade assembly provides 30 and the present main cooling circuit 60 a new advantageous solution for a cooling problem with a blade arrangement recognized by the present inventors 30 ,

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including the creation and use of any devices or systems, and the performance of any incorporated methods belong. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

There are created blade assemblies. The bucket assembly includes a shank and an airfoil positioned radially outward of the shank. The bucket assembly further includes a main cooling circuit formed in the airfoil and shaft, the main cooling circuit having seven passages, each of the seven passages being fluidly connected to an adjacent one of the seven passages. A maximum rotation number in each of the seven passes is less than or equal to about 0.4.

LIST OF REFERENCE NUMBERS

10

Gas Turbine System

12

compressor section

14

combustor section

16

turbine section

18

wave

19

inlet section

20

outlet

21

Vanes of the first stage

22

Rotor blades of the first stage

23

Guide vanes of the second stage

24

Blades of the second stage

25

Third stage vanes

26

Blades of the third stage

28

Hot gas flow

30

Blade assembly

32

shaft

34

airfoil

36

platform

38

dovetail

42

pressure side

44

suction

46

leading edge

48

trailing edge

52

top

54

root

60

Main cooling circuit

61

First cooling passage

62

Second cooling passage

63

Third cooling passage

64

Fourth cooling passage

65

Fifth cooling passage

66

Sixth cooling passage

67

Seventh cooling passage

68

First inlet section

69

Second inlet section

70

Front cooling circuit

71

First cooling passage

72

Second cooling passage

80

Rear cooling circuit

81

First cooling passage

90

outlet

a

Cross-section length of a passage

b

Section width of a passage

Claims (10)

Blade assembly comprising: a shaft; an airfoil positioned radially outward of the shaft; a main cooling circuit defined in the airfoil and the shaft, the main cooling circuit having seven passages, each of the seven passages being fluidly connected to an adjacent one of the seven passages, wherein a maximum rotation number in each of the seven passes is less than or equal to about 0.4. The blade assembly of claim 1, wherein the rotation number in each of the seven passes is in a range between about 0.01 and about 0.35. The blade assembly of claim 1, wherein a first passage of the seven passages has a maximum cross-sectional area in the airfoil that is smaller by a size in a range between about 10% and about 20% than a maximum cross-sectional area of another passage of the seven passages in the airfoil , The blade assembly of claim 1, further comprising an outlet passage defined in the airfoil and in fluid communication with a seventh passage of the seven passages. The blade assembly of claim 1, further comprising a front cooling circuit and a rear cooling circuit defined respectively in the airfoil and the shaft. The blade assembly of claim 1, wherein the main cooling circuit consists of seven passes. The blade assembly of claim 1, further comprising a platform positioned radially between the shaft and the airfoil. The blade assembly of claim 1, wherein the airfoil has outer surfaces defining a pressure side and a suction side extending between a leading edge and a trailing edge. The blade assembly of claim 1, further comprising a dovetail, wherein the dovetail defines a root and wherein the main cooling circuit preferably extends to the root. A turbomachine comprising: an inlet section; an outlet section; a compressor section; a combustion chamber section; and a turbine section, the turbine section having a plurality of blade assemblies, each of the plurality of blade assemblies any one of the preceding claims.

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