WO2018153796A1

WO2018153796A1 - A turbomachine blade or vane having a cooling channel with a criss-cross arrangement of pins

Info

Publication number: WO2018153796A1
Application number: PCT/EP2018/053952
Authority: WO
Inventors: Andrea Viano
Original assignee: Siemens Aktiengesellschaft
Priority date: 2017-02-24
Filing date: 2018-02-16
Publication date: 2018-08-30

Abstract

A turbomachine component having an aerofoil, such as a blade or a vane, is presented. The turbomachine component includes an aerofoil cavity bordered by a suction side wall and a pressure side wall. A cooling channel, for flow of cooling fluid, defined by an internal wall (60) of the cooling channel extends inside the aerofoil cavity. Within the cooling channel, each of the first and the second pins (71, 81) is connected to opposite sides of the internal wall and extends from wall-to-wall. The first and the second pins are non-contiguously and alternatively arranged within the cooling channel and are disposed at an angle, with respect to each other, thereby forming a criss-cross arrangement of the pins. The wall-to-wall extending pins provide greater surface area for heat conduction, generate turbulence in the cooling fluid, and redirect heat from one part of the turbomachine component to another part of the turbomachine component by conduction.

Description

A turbomachine blade or vane having a cooling channel with a criss-cross arrangement of pins

The present invention relates to gas turbines, and more particularly to blades or vanes of gas turbines with cooling channels . Cooling of gas turbine components, such as a turbine blade or a vane is a major challenge and an area of interest in turbine technology. A common technique for cooling a turbine blade/vane, i.e. blade and/or vane, is to have one or more internal passages, referred to as cooling channels or cooling passages, within the blade/vane and to flow a cooling fluid, such as cooling air through these internal cooling channel. Surfaces of such cooling channel or channels are often lined with turbulators to enhance the heat transfer into the cooling air from the blade/vane internal surfaces forming surfaces of the cooling channel i.e. forming internal wall of the cooling channel. Often a series of rib turbulators or pin-fin turbulators are arranged along the flow path of the cooling fluid within the cooling channel. The turbulators induce turbulence in the cooling fluid and thereby increase the efficiency of the heat transfer.

Although use of turbulators increases the efficiency of heat transfer, there is a scope of further improvement and

therefore constant efforts are put in by engineers, designers and researchers to modify the turbulator structures,

arrangements, orientation and positioning within the cooling channel to increase, enhance or maintain the efficiency of cooling. Furthermore, other types of techniques, instead of or in addition to using the turbulators, are also further developed continuously. The concerted efforts have resulted in various cooling techniques that employ ribs, dimples, impingement cooling, pin fins, matrix cooling, vortex

cooling, and so on and so forth. However, a scope of

improvement in existing cooling techniques or introduction of alternate techniques for cooling of a turbomachine component having cooling channels constantly exists.

Thus the object of the present disclosure is to provide a technique for cooling a turbomachine component having an aerofoil and a cooling channel. It is desirable that the present technique provides high heat transfer coefficients within the cooling channel thereby resulting in efficient cooling of the turbomachine component having the aerofoil.

The above objects are achieved by a turbomachine component according to claim 1 of the present technique. Advantageous embodiments of the present technique are provided in

dependent claims. Features of claim 1 can be combined with features of dependent claims, and features of dependent claims can be combined together.

The present technique presents a turbomachine component which has an aerofoil. An example of such turbomachine component is a blade or a vane for a turbomachine or a gas turbine engine. The aerofoil of the turbomachine component includes a suction side wall and a pressure side wall. The side walls, namely the suction side wall and the pressure side wall bordering an aerofoil cavity. The turbomachine component also has at least one cooling channel that extends inside at least a part of the aerofoil cavity. The cooling channel is defined within the aerofoil cavity by an internal wall of the cooling channel. The cooling channel is for flow of a cooling fluid, when present. The cooling channel has an inlet that receives the cooling fluid which then flows through the cooling channel. Furthermore, the turbomachine component includes, arranged inside the cooling channel, at least a first row of pins and a second row of pins. The first row of pins has a plurality of first pins. The pins, hereinafter also referred to as the pin fins, of the first row are elongated cylindrical structures having

circular or polyhedral cross-sections, for example a rod like structure. Each of the first pins has two ends that are longitudinally spaced apart by an elongated body of the first pin, like a rod having two ends and the curved surface of the rod forming the elongated body. One end of each first pin is connected i.e. is physically continuous with or affixed to the internal wall of the cooling channel and the other end of each first pin is also connected i.e. is physically

continuous with or affixed to the internal wall of the cooling channel on opposite side. Each first pin extends longitudinally within the cooling channel. Simply understood, the two ends of each first pin are in physical contact with the internal wall of the cooling channel on opposite sides and the elongated body of the first pin is hanging in or disposed in space of the cooling channel through which the cooling fluid flows, when present, for example if the cooling channel is tubular with circular cross-section each of the first pin is oriented like a chord of the circular cross- section, and different first pins are aligned or located lengthwise along the cooling channel thereby forming the first row.

Similarly, the second row of pins has a plurality of second pins. The pins, hereinafter also referred to as the pin fins, of the second row are elongated cylindrical structures having circular or polyhedral cross-sections, for example a rod like structure. Each of the second pins has two ends that are longitudinally spaced apart by an elongated body of the second pin, like a rod having two ends and the curved surface of the rod forming the elongated body. One end of each second pin is connected i.e. is physically continuous with or affixed to the internal wall of the cooling channel and the other end of each second pin is also connected i.e. is physically continuous with or affixed to the internal wall of the cooling channel on opposite side. Each second pin extends longitudinally within the cooling channel. Simply understood, the two ends of each second pin are in physical contact with the internal wall of the cooling channel on opposite sides and the elongated body of the second pin is hanging in or disposed in space of the cooling channel through which the cooling fluid flows, when present, for aforementioned example of the tubular cooling channel with circular cross-section each of the second pin is oriented like a chord of the circular cross-section, and different second pins are aligned or located lengthwise along the cooling channel thereby forming the second row.

Within the cooling channel, the first pins and the second pins, with respect to each other, are non-contiguously and alternatively arranged i.e. when moving into the cooling channel in the flow direction for the cooling fluid from say the inlet of the cooling channel, first comes the first pin then the second pin and then the first pin and then the second pin and so on and so forth. The first pins and the second pins are non-contiguous i.e. the first pins and the second pins are not in physical contact of each other, rather are physically separated from one another. Furthermore, the first pins and the second pins are disposed at an angle with respect to each other i.e. when viewing into the cooling channel in the flow direction for the cooling fluid from say the inlet of the cooling channel, the first pins and the second pins appear to form a criss-cross pattern within the cooling channel. Simply put, the first row and the second row are mutually interspersed and aligned at an angle with respect to each other. In an embodiment of the turbomachine component, the alternatively arranged first pins and second pins are equidistantly spaced apart from one another, the flow direction for the cooling fluid within the cooling channel . Each of the first pins have a centreline and a length and each of the second pins have a centreline and a length. The centrelines cross each other at a location within at least 10% of one of the lengths from the internal wall. The present technique, therefore, provides an alternative cooling arrangement with respect to the conventionally known cooling arrangements. The criss-cross arrangement of the first and the second pins increases turbulence and local velocity in the cooling fluid, when flowing in the cooling channel. The flow, i.e. the flowing cooling fluid, is

redirected around the pins, i.e. the first and the second pins, and is diverted and redirected continuously within the cooling channel due to the different orientation of the first pins and the second pins encountered by the flow within the cooling channel. Furthermore, since the pins, i.e. the first and the second pins, extend wall-to-wall within the internal wall of the cooling channel, the surface area of the internal wall of the cooling channel is greater than the

conventionally known turbulator structures that emanate outward from the walls of the cooling channel but do not extend up to the other side of the wall. Therefore, by the first and the second pins of the present technique, the surface area within the cooling channel available for

conductive heat transfer to the cooling fluid is increased.

Moreover, due to presence of the wall-to-wall extending first and the second pins resulting in more surface area available for conductive cooling, a desired cooling effectiveness is reached with less mass flow. This is positive for the engine performance because less air is used to cool the turbomachine component .

The angle, i.e. the smaller angle out of the two adjacent angles that the first and the second pins form when viewed from a direction mutually perpendicular to directions of longitudinal extension of the first and the second pins, at which the first pins and the second pins are disposed with respect to each other is between 20 degree and 90 degree, and more particularly between 45 degree and 90 degree. When viewing into the cooling channel in the flow direction for the cooling fluid, the criss-cross pattern that the first pins and the second pins appear to form within the cooling channel is ^ΛΧ' shaped wherein out of the two arms of the ^ΛΧ' one is the first pin, or the first row, and the other is the second pin, or the second row.

In an embodiment of the turbomachine component, the angle is 90 degree i.e. the first pins and the second pins are

arranged orthogonally. When viewing into the cooling channel in the flow direction for the cooling fluid, the criss-cross pattern that the first pins and the second pins appear to form within the cooling channel is cross shaped. In another embodiment of the turbomachine component, the internal wall of the cooling channel has a plurality of sections arranged circumferentially with respect to the cooling channel, i.e. circumferentially to the flow direction of the cooling fluid, when present and flowing. The plurality of sections includes at least one pressure side wall section, at least one suction side wall section and at least one rib section. The pressure side wall section is formed by a part of the pressure side wall i.e. the pressure side wall of the aerofoil has two faces - one face forming part of the

external surface of the aerofoil and which is exposed to the hot gas flow during operation of the engine and other face forming a part of the internal wall of the cooling channel. Thus, the pressure side wall section of the internal wall of the cooling channel is the section of the internal wall of the cooling channel that is towards the or in vicinity of the pressure side of the aerofoil. The suction side wall section is formed by a part of the suction side wall i.e. the suction side wall of the aerofoil has two faces - one face forming part of the external surface of the aerofoil and which is exposed to the hot gas flow during operation of the engine and other face forming a part of the internal wall of the cooling channel. Thus, the suction side wall section of the internal wall of the cooling channel is the section of the internal wall of the cooling channel that is towards the or in vicinity of the suction side of the aerofoil. The rib section is formed by a part of surface of a rib that is present in the aerofoil cavity and generally extending between the pressure side and suction side walls. Thus, the rib section of the internal wall of the cooling channel is the section of the internal wall of the cooling channel that is formed by the rib of the aerofoil. The internal wall of the cooling channel may include two opposing rib sections formed by two separate ribs of the aerofoil. In one embodiment of the turbomachine component the first pins of the first row and the second pins of the second row extend between the pressure side wall section and the suction side wall section of the internal wall of the cooling

channel. Thus the pins, i.e. the first and the second pins, conduct heat from the pressure side wall and the suction side wall, hereinafter also referred to as the side walls, into the cooling channel whereby the heat can be transferred to the cooling fluid.

In another embodiment of the turbomachine component the first pins of the first row or the second pins of the second row extend between the pressure side wall section and the suction side wall section of the internal wall of the cooling channel whereas the other of the first pins of the first row and the second pins of the second row extend between the rib section and the pressure side or the suction side wall section of the internal wall of the cooling channel. Thus the one of the pins, i.e. the first pins or the second pins, conduct heat from the pressure side wall or the suction side wall into the cooling channel whereby the heat can be transferred to the cooling fluid, whereas the other of the pins, i.e. the first pins or the second pins, conduct heat from the pressure side wall or the suction side wall into the cooling channel and further to the rib section whereby the heat can be

transferred to the cooling fluid. Since the rib section is relatively cooler than the side walls, the coefficient of heat transfer for the pins connecting the rib section with any of the side walls is increased, and thereby increasing the efficiency of cooling achieved by the cooling fluid.

Alternatively, in another embodiment of the turbomachine component having two rib sections in the internal wall of the cooling channel, the first pins of the first row or the second pins of the second row extend between the pressure side wall section and the suction side wall section of the internal wall of the cooling channel whereas the other of the first pins of the first row and the second pins of the second row extend between the two opposing rib sections of the internal wall of the cooling channel. Thus the one of the pins, i.e. the first pins or the second pins, conduct heat from the pressure side wall or the suction side wall into the cooling channel whereby the heat can be transferred to the cooling fluid, whereas the other of the pins, i.e. the first pins or the second pins, being connected to two opposing rib sections redistributes heat between the two ribs and into the cooling channel whereby the heat can be transferred to the cooling fluid. The redistribution of heat or the thermal conduction between the two ribs enables heat flow from two adjacent sections of the cooling channel divided by an intermediately positioned rib or between two separate cooling channels divided by an intermediately positioned rib, and thereby increasing the efficiency of cooling achieved by the cooling fluid.

In yet another embodiment of the turbomachine component, the pressure side wall section and the suction side wall section meet each other forming an angular section of the internal wall of the cooling channel. The angular section may be in form of an edge or a rounded edge. The angular section of the internal wall is the section of the internal wall formed towards or in vicinity of a leading edge of the aerofoil or is the section of the internal wall formed towards or in vicinity of a trailing edge of the aerofoil. The first pins of the first row or the second pins of the second row extend between angular section and to the rib section of the

internal wall of the cooling channel whereas the other of the first pins of the first row and the second pins of the second row extend between the pressure side wall section and the suction side wall section of the internal wall of the cooling channel. This enables conduction of heat from the angular section into the pins and subsequently towards the rib section. This is especially beneficial where the angular section of the internal wall is positioned towards the leading edge, as the connection of the angular section to the ribs via the pins increases the coefficient of heat transfer owing to the significant difference in temperature of the leading edge and the rib. In another embodiment of the turbomachine component, the first pins and/or the second pins are formed as one-part extension of the internal wall of the cooling channel. This may be achieved by forming the pins integrally with the internal wall of the cooling channel, for example by casting or additive manufacturing techniques, and enables better heat conduction between the internal wall of the cooling channel and the pins. Alternatively, in another embodiment of the turbomachine component, the first pins and/or the second pins are formed as a fixture attached to the internal wall of the cooling channel.

The above mentioned attributes and other features and

advantages of the present technique and the manner of

attaining them will become more apparent and the present technique itself will be better understood by reference to the following description of embodiments of the present technique taken in conjunction with the accompanying

drawings, wherein:

FIG 1 shows part of a turbine engine in a sectional view and in which a turbomachine component of the present technique is incorporated;

FIG 2 schematically illustrates a perspective view of an exemplary embodiment of the turbomachine component with an aerofoil;

FIG 3 schematically illustrates perspective view of an exemplary embodiment of the turbomachine component with a part of the turbomachine component removed to depict cooling channels inside the turbomachine component ;

FIG 4 schematically illustrates a cross-section of

exemplary embodiment of the aerofoil normal

longitudinal axis of the aerofoil; schematically illustrates a perspective view of a part of the turbomachine component depicting an exemplary arrangement, according to the present technique, of first pins and second pins in a part of the cooling channel of the turbomachine

component ; schematically illustrates a perspective view of a part of the turbomachine component depicting an exemplary arrangement, according to the present technique, of the first pins and the second pins in another part of the cooling channel of the

turbomachine component; schematically illustrates a perspective view of a part of the turbomachine component depicting an exemplary arrangement, according to the present technique, of the first pins and the second pins in yet another part of the cooling channel of the turbomachine component; ; schematically illustrates a part of the cross- section of the aerofoil shown in FIG 4 and depicts an exemplary positioning of the first pins and the second pins arranged in the cooling channel; schematically illustrates the part of FIG 8 and depicts another exemplary positioning of the first pins and the second pins arranged in the cooling channel ; schematically illustrates the part of FIG 8 and depicts yet another exemplary positioning of the first pins and the second pins arranged in the cooling channel; schematically illustrates another part of the cross-section of the aerofoil, as compared to the part shown in FIG 4, and depicts an exemplary positioning of the first pins and the second pins arranged in the cooling channel; schematically illustrates the part of FIG 11 and depicts another exemplary positioning of the first pins and the second pins arranged in the cooling channel ; schematically illustrates cross sectional view of a part of a cooling channel with the second pins angled in relation to a flow direction of the fluid, in accordance with the present technique; shows a perspective view of an aerofoil with dashed hidden lines depicting a number of cooling channels extending through the aerofoil and comprising an array of the first pins and the second pins arranged in the cooling channel; shows another perspective view on the pressure side of the aerofoil of FIG 14; and is a view radially inwardly on a cooling channel of a blade's aerofoil and showing rows of first pins and second pins that are twisted.

Hereinafter, above-mentioned and other features of the present technique are described in details. Various

embodiments are described with reference to the drawing, wherein like reference numerals are used to refer to like elements throughout. In the following description, for the purpose of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be noted that the illustrated embodiments are intended to explain, and not to limit the invention. It may be evident that such embodiments may be practiced without these specific details. It may be noted that in the present disclosure, the terms "first", "second", etc. are used herein only to facilitate discussion, and carry no particular temporal or chronological significance unless otherwise indicated.

FIG. 1 shows an example of a gas turbine engine 10 in a sectional view. The gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor or compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20. The gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10. The shaft 22 drivingly connects the turbine section 18 to the

compressor section 14.

In operation of the gas turbine engine 10, air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16. The burner section 16 comprises a longitudinal axis 35 of the burner, a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28. The combustion chambers 28 and the burners 30 are located inside the burner plenum 26. The compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17. This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for

channelling the combustion gases to the turbine 18. The turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22. In the present example, two discs 36 each carry an annular array of turbine blades 38. However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, guiding vanes 40, which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.

The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotates the shaft 22. The guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.

The turbine section 18 drives the compressor section 14. The compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48. The rotor blade stages 48 comprise a rotor disc supporting an annular array of blades. The compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48. The guide vane stages include an annular array of

radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operational conditions.

The casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14. A radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48. The present technique is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present technique is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.

The terms axial, radial and circumferential are made with reference to the rotational axis 20 of the engine, unless otherwise stated.

It may be noted that the present technique has been explained in details with respect to an embodiment of a turbine blade, however, it must be appreciated that the present technique is equally applicable and implemented similarly with respect to a turbine vane or any other turbomachine component being cooled by a cooling channel that extends within the

turbomachine component. FIG 2 schematically illustrates a turbomachine component 1 having an aerofoil 5, for example the turbine blade 38 or the vane 40 of FIG 1. FIG 3 shows the turbomachine component 1 with a part of the turbomachine component 1 removed to depict cooling channels 6 inside the turbomachine component 1.

Hereinafter the turbomachine component 1 has also been referred to as the blade 1. In the blade 1, the aerofoil 5 extends from a platform 68 in a radial direction 97 which also represents a longitudinal axis 97 for the blade. From another side of the platform 68 emanates a root 69 or a fixing part 69 that is used to attach the blade 1 to the turbine disc 38 (shown in FIG 1) . It may be noted that in some other embodiments of the turbomachine component 1, the root 69 may not be present, and the turbomachine component 1 may be an integrally fabricated part of a larger structure (not shown) such as a stator disc in the turbine section 16 of the engine 10, as shown in FIG 1.

The aerofoil 5 includes a suction side wall 2, also called suction side 2, and a pressure side wall 3, also called pressure side 3. The side walls 2 and 3 meet at a trailing edge 92 on one end and a leading edge 91 on another end. The aerofoil 5 has a tip end opposite to the platform 68. The aerofoil 5 may be connected to a shroud (not shown) at the tip end of the aerofoil 5. In some other embodiments the aerofoil 5 may be connected to a tip platform (not shown) instead of the shroud. The shroud and the tip platform are commonly referred to a tip (not shown) of the turbomachine component 1. The aerofoil 5 may also include a shroud (not shown) at the tip end of the aerofoil 5. The side walls 2 and 3 of the aerofoil 5 act as boundary for an aerofoil cavity 4.

Referring to FIG 3, an exemplary embodiment of the

turbomachine component 1, i.e. the blade 1 has been explained hereinafter. The blade 1 has at least one cooling channel 6 that extends inside at least a part of the aerofoil cavity 4. A cooling fluid, such as cooling air, flows through the cooling channel 6 and direction of flow of the cooling fluid has been represented by arrows marked with reference numeral 7. The cooling channel 6 has an inlet 66 that receives the cooling fluid which then flows through the cooling channel 6. The cooling channel 6 usually has a serpentine path though the aerofoil cavity 4. Within the aerofoil cavity 4, are ribs 75 that divide the aerofoil cavity 4 to accommodate multiple separate cooling channels 6 as shown in FIG 3 or that divide a given cooling channel 6 into separate sections formed in a serpentine manner. The cooling fluid after flowing through the cooling channel 6 exits the cooling channel 6 for example by holes 95 that fluidly connect the cooling channel 6 to an outside of the aerofoil 5. The holes 95 may be present at any region of the aerofoil 5 for example at the trailing edge 92 or a leading edge 91. FIG 4 shows a cross-section of aerofoil 5 depicting two ribs 75 and 85 that form different cooling channels 6 and/or that create separate sections of a given cooling channel 6. FIG 4 depicts three different sections of the cooling channel 6 or three different cooling channels 6, or a combination of both and hereinafter referred to as the first cooling channel 6, the second cooling channel 6 and the third cooling channel 6 ordered from the leading edge 91 to the trailing edge 92. As can be seen in FIG 4 the cooling channel 6 is defined by an internal wall 60 of the cooling channel 6. The internal wall 60 of the cooling channel 6 has different segments or

sections, arranged circumferentially with respect to the cooling channel i.e. normally to the flow direction 7 (as shown in FIG 3) of the cooling fluid within the cooling channel 6. The sections of the internal wall 60, as shown in the FIG 4, are a pressure side wall section 61, a suction side wall section 62 and a rib section 63 or 64.

The pressure side wall section 61 is the section of the internal wall 60 of the cooling channel 6 that is formed by a part of the pressure side wall 3 i.e. the section of the internal wall 60 adjacent to or in vicinity of the pressure side wall 3 of the aerofoil 5. The suction side wall section 62 is the section of the internal wall 60 of the cooling channel 6 that is formed by a part of the suction side wall 2 i.e. the section of the internal wall 60 adjacent to or in vicinity of the suction side wall 2 of the aerofoil 5. The rib section 63, 64 is the section of the internal wall 60 of the cooling channel 6 that is formed by a part of the rib 75, 85 or is adjacent to or in vicinity of the rib 75, 85 of the aerofoil 5. As shown in FIG 4, the first and the third cooling channel 6 have one rib section 63 whereas the second cooling channel 6 has two rib sections 63 and 64.

Furthermore, as shown in the first and the third cooling channel 6, the pressure side wall section 61 and the suction side wall section 62 meet or converge together forming an angular section 65 of the internal wall 60 of the cooling channel 6. The angular section 65 may either be located at the leading edge 91 as for first cooling channel 6 or at the trailing edge 92 as for the third cooling channel 6.

Further details of the turbomachine component 1, i.e. the blade 1, according to the present technique are explained with reference to FIGs 5 to 12 in combination with FIG 4. The turbomachine component 1 includes at least a first row 70 of pins and a second row 80 of pins. FIG 5 shows arrangement of the first row 70 of pins and the second row 80 of pins in the first cooling channel 6, as shown in FIG 4, whereas FIGs 6 and 7 show arrangement of the first row 70 of pins and the second row 80 of pins in the second and the third cooling channel 6, as is referred to in FIG 4. As shown in FIGs 5 to 7, the first row 70 of pins includes a plurality of first pins 71. One end of each first pin 71 is connected to the internal wall 60 of the cooling channel 6 and another end of each first pin 71 is connected to the internal wall 60 of the cooling channel 6 on opposite side such that each first pin 71 extends within the cooling channel 6 from the internal wall 60 to the internal wall 60 oriented substantially perpendicular to flow direction of the cooling fluid, or simply put wall-to-wall. Similarly, the second row 80 of pins includes a plurality of second pins 81. One end of each second pin 81 is connected to the internal wall 60 of the cooling channel 6 and another end of each second pin 81 is connected to the internal wall 60 of the cooling channel 6 on opposite side such that each second pin 81 extends within the cooling channel 6 from the internal wall 60 to the internal wall 60 oriented substantially perpendicular to flow direction of the cooling fluid, simply put wall-to-wall. As shown in FIGs 5 to 7, the first pins 71 and the second pins 81, with respect to each other, are non- contiguously and alternatively arranged within the cooling channel 6, i.e. the any given first pin 71 is separated from its neighbouring second pin 81 or second pins 81 by a

distance. In an exemplary embodiment of the turbomachine component 1, and as depicted in FIGs 5 to 7, the alternatively arranged first pins 71 and second pins 81 are equidistantly spaced apart from one another.

The first pins 71 and the second pins 81, with respect to each other, and are disposed at an angle 99 (shown in FIGs 8 and 9, not shown in FIGs 5 to 7) . As a result of the mutual angular disposition, the first pins 71 and the second pins 81 form a criss-cross pattern or arrangement within the cooling channel 6. Further details of orientation and positioning of the first pins 71 and the second pins 81 are provided, hereinafter later, with respect to FIGs 8 to 12.

The flow 7 of cooling fluid, i.e. the flowing cooling fluid 7, when present and flowing, flows about the first pins 71 and the second pins 81, encountering each in series, and thus the flow 7 of the cooling fluid is twisted and turned within the cooling channel 6. In one embodiment of the component 1, one or both of the first pins 71 and the second pins 81 are orthogonal to the flow 7 or inclined at an angle 100 (as shown in FIG 13) different than 90 degree. In particular the angle 100 may be between 20 degree and 160 degree. At least a part of the flowing cooling fluid 7 while flowing within the cooling channel 6 flows in contact with the first pins 71 and the second pins 81, hereinafter also referred collectively as the pins 71, 81. The external shape of the pins 71, 81 is such that turbulence or swirl is introduced in the flowing cooling fluid 7 as a result of contacting or flowing around the pins 71, 81. The shape and dimensions of the pins 71, 81 are such that turbulence is generated, for example a vortex or vortices are generated in the flowing cooling fluid 7 inside the cooling channel 6. As a result of the swirl generation the cooling effect of the cooling fluid 7 is enhanced. Furthermore, the twisting and turning of the flow 7 of the cooling fluid resultant from encountering the pins 71, 81 directs the flow towards the internal wall 60 of the cooling channel 6 and thus cooling the internal wall 60. The surface area within the cooling channel 6 available for conductive heat loss to the cooling fluid is provided by the internal wall 60 and a surface of the wall-to-wall extending pins 71, 81.

Hereinafter, referring to FIGs 8 to 12, further details of orientation and positioning of the pins 71, 81 are provided. FIGs 8, 9 and 10 represent the second cooling channel 6, as referred to in FIG 4, and depict possibilities of different orientations and positions of the pins 71, 81; whereas FIGs 11 and 12 represent the first cooling channel 6, as referred to in FIG 4, and depict possibilities of different

orientations and positions of the pins 71, 81. It may be noted that the hereinafter mentioned orientations and

positions of the pins 71, 81 with respect to the first cooling channel 6 and the second cooling channel 6 are for exemplary purposes only. As may be appreciated by one skilled in the art, the orientations and positions of the pins 71, 81 explained hereinafter with reference to FIGs 8 to 12 are also applicable for other shapes of the cooling channel 6. The criss-cross arrangement of the pins 71, 81, when viewed in a direction mutually perpendicular to the pins 71, 81 is achieved by the mutual angular disposition of the pins 71 81. The angle 99 may be between 20 degrees and 90 degrees, as shown in FIG 9 that shows the angle 99 to be less than 90 degrees and in FIG 8 that shows the angle 99 to be about or equal to 90 degrees resulting into the criss-cross

arrangement having substantially orthogonal mutual

disposition. The angle 99 is the smaller angle out of the two adjacent angles that the first and the second pins 71, 81 form when viewed from a direction mutually perpendicular to directions of longitudinal extension of the first and the second pins 71, 81.

As shown in FIGs 8, 9 and 10, and also in FIG 4, the cooling channel 6 i.e. for example the second cooling channel 6 of

FIG 4, is defined by the internal wall 60 having the pressure side wall section 61, the suction side wall section 62 opposite to the pressure side wall section 61, the rib sections 63 along the rib 75 and the rib section 64 along the rib 85. The different positioning of the pins 71, 81 are explained with respect to the sections 61, 62, 63, 64 in the FIGs 8, 9 and 10. It may be noted that although in FIGs 8 to 10 only one first pin 71 and one second pin 81 are depicted, because FIGs 8 to 10 are cross-sectional views, the other first pins 71 of the first row 70 and the other second pins 81 of the second row 80 are similarly positioned.

As shown in FIG 8, the first pins 71 extends between, or is connected to or contiguous with, the two rib section 63, 64, whereas the second pin 81 extends between, i.e. is connected to or contiguous with, the pressure side wall section 61 and the suction side wall section 62 of the internal wall 60 of the cooling channel 6.

As shown in FIG 9, the first pins 71 extends between, or is connected to or contiguous with, the two rib section 63, 64, whereas the second pin 81 extends between, i.e. is connected to or contiguous with, the pressure side wall section 61 and the rib section 63 of the internal wall 60 of the cooling channel 6. In another embodiment (not shown), the first pins 71 extends between, or is connected to or contiguous with, the two rib section 63, 64, whereas the second pin 81 extends between, i.e. is connected to or contiguous with, the suction side wall section 62 and the rib section 63. In yet another embodiment (not shown) , the first pins 71 extends between the two rib section 63, 64, whereas the second pin 81 extends between the suction side wall section 62 and the rib section 64. In a further embodiment (not shown), the first pins 71 extends between the two rib section 63, 64, whereas the second pin 81 extends between the pressure side wall section 61 and the rib section 64.

As shown in FIG 10, the first pins 71 extends between, i.e. is connected to or contiguous with, the rib section 63 and the suction side wall section 62, whereas the second pin 81 extends between, i.e. is connected to or contiguous with, the pressure side wall section 61 and the rib section 63 of the internal wall 60 of the cooling channel 6. In another embodiment (not shown) , the first pins 71 extends between, the rib section 64 and the suction side wall section 62, whereas the second pin 81 extends between the pressure side wall section 61 and the rib section 64 of the internal wall 60 of the cooling channel 6.

As shown in FIGs 11 and 12, and also in FIG 4, the cooling channel 6 i.e. for example the first cooling channel 6 of FIG 4, is defined by the internal wall 60 having the pressure side wall section 61, the suction side wall section 62 opposite to the pressure side wall section 61, the rib sections 63 along the rib 75. The suction side wall section 62 converges with or is formed with the pressure side wall section 61 via an intermediate section referred to as the angular section 65. The different positioning of the pins 71, 81 are explained with respect to the sections 61, 62, 63, 65 in the FIGs 11 and 12. It may be noted that although in FIGs 11 and 12 only one first pin 71 and one second pin 81 are depicted, because FIGs 11 and 12 are cross-sectional views, the other first pins 71 of the first row 70 and the other second pins 81 of the second row 80 are similarly positioned.

As shown in FIG 11, the first pins 71 extends between, or is connected to or contiguous with, the angular section 65 and the rib section 63, whereas the second pin 81 extends

between, i.e. is connected to or contiguous with, the

pressure side wall section 61 and the suction side wall section 62 of the internal wall 60 of the cooling channel 6. Alternatively, as shown in FIG 12, the first pins 71 and the second pins 81 are both connected to the rib section 63 and to the internal wall 60 but offset from center of the angular section 65 or, simply put connected to sides of the angular section 65, and thereby leaving a central space within the angular section 65 free. The holes 95, shown in FIG 3, may thus be positioned within the central space of the angular section 65 for allowing exit of the cooling fluid, and thereby achieving film cooling, at the leading edge 91 of the aerofoil 5. It may be noted that although in FIGs 11 and 12, the first cooling channel 6, as referred to in FIG 4, is depicted, the same orientation and positioning is applicable for the third cooling channel 6 of FIG 4 in which the angular section 65 is at the trailing edge 92.

As a result of connections established between different sections 61, 62, 63, 64, 65 of the internal wall 60 by the pins 71, 81, a conductive path for heat transfer between the different sections 61, 62, 63, 64, 65 through the body of the pins 71, 81 is established, which helps in heat distribution within the aerofoil 5 and thus reduces development of high localized thermal stresses. As an example, in embodiment of the turbomachine component 1 shown in FIG 11, conductive heat from is established from the leading edge 91 to the rib 75 which is relatively cooler. Furthermore since the pins 71, 81 connected to the sections 61, 62, 63, 64, 65 of the internal wall 60 are disposed within the cooling channel 6, the heat conducted from the sections 61, 62, 63, 64, 65 of the

internal wall 60, and thereby from their adjoining parts of the aerofoil 5, is available for conductive transfer when the cooling fluid flows through the cooling channel 6 and around the pins 71, 81. The pins 71, 81 may be formed as one-part extension of the internal wall 60 of the cooling channel 6 for example during casting or forging of the turbomachine component 1 along with the cooling channel 6, or are formed as a fixture attached to the internal wall 60 of the cooling channel 6.

FIG 14 shows a perspective view of the aerofoil 5 as

described hereinbefore. The dashed or hidden lines depict a number of the cooling channels 6 which extend through the aerofoil 5. The channels 6 comprise an array of the first pins 70 and the second pins 80 arranged in the cooling channel 6 as hereinbefore described. FIG 15 shows another perspective view on the pressure side 3 of the aerofoil 5 of FIG 14. In this example, the aerofoil 5 is often used in a part of the gas turbine which is cooler than the previously described examples. It is possible to describe this aerofoil as a ^Asolid' as opposed to a substantially ^,hollow' aerofoil. The aerofoil 5 has an overall cross-sectional area A and the cooling channels have a total cross sectional area C. It may be said that for a solid blade the ratio of C/A is less than 40% and may be less than 20%. The present arrangement of pins 70 and 80 is implemented here to provide the advantages described herein and in addition create a blockage to reduce the mass flow of coolant through the cooling channels 6 while increasing the heat transfer coefficient. This is

particularly, but not exclusively, desirable where the direction of flow of the coolant is in only one radial direction through all the cooling channels 6 in the aerofoil 6.

Any one pin 71, 81 may occupy 30-70% of the cross-sectional area of the cooling channel 6 in a direction perpendicular to the direction of flow of the coolant to provide good heat transfer and reduce the mass flow of the coolant. Thereby cooling efficiency is greatly improved. Preferably, any one pit 71, 81 may occupy 40-50% of the cross-sectional area of the cooling channel 6.

In general, but referring to FIG. 8, the present arrangement of first pins 71 and second pins 81 is now described with more detail and in particular the relationship of the

location where the first and second pins cross each other, crossing point P. Each of the first pins 71 have a

centreline CL1 and a length LI and each of the second pins 81 have a centreline CL2 and a length L2. Importantly, for any two consecutive first and second pins 71, 81, i.e. adjacent or neighbouring or pair of pins, the centrelines CL1, CL2 cross each other at a location within at least 10% of one of the lengths LI, L2 from the internal wall 60. What is meant by ^,cross each other' is with respect to the view shown in

Fig. 8 or which is perpendicular to the main or longitudinal extent of the cooling channel 6 or perpendicular to the direction of the main flow of coolant passing through the channel 6. Thus the centrelines cross each other viewed in plan or in Fig. 8 viewed radially. Furthermore, ^,cross each other' is not intended to mean that the centrelines

physically intersect one another. In defining that the first and second pins 71, 81 cross or cross over each other, the wake or vortex created by one pin impinges on the downstream pin. This arrangement therefore enhances turbulence and mixing of the coolant flow in the channel 6 and also convection of heat in the downstream pins. Preferably, the centrelines of first and second pins 71, 81 cross each other at a location within 20% of one of the lengths LI, L2 from the internal wall 60 and even more preferably within 30% of one of the lengths LI, L2 from the internal wall 60. This means that the centrelines cross each other in a location that is at least 10% (20% or 30%) from both ends of say pin 71 or indeed pin 81.

In the embodiments shown and described herein any two

consecutive pins 71, 81 form a cross at an angle when viewed along the longitudinal direction of the cooling channel 6.

It can be said that the cross shape distinguishes four ^pen' sectors SI, S2, S3 and S4 as shown in FIG.8. In other words there are four areas that have a clear line-of-sight and that cooling air can pass in the longitudinal direction of the channel 6.

The pins are preferably cylindrical having a generally circular cross-section, but other cross-sectional shapes are possible. As seen in FIG.7 the pins have a width W in a direction perpendicular to the longitudinal direction of the channel 6. Consecutive first pins 71 and second pins 81 are separated or spaced, centreline to centreline, a distance X in the longitudinal direction of the channel 6. For the optimal formation of turbulence and cooling the spacing X is between 2W and 10W and preferably approximately 6W.

Furthermore, consecutive first and second pins 71, 81 do not need to be equally spaced and instead where more cooling is required (i.e. the aerofoil experiences higher temperatures) the spacing can be reduced to increase cooling. It is possible to tailor the spacing X of consecutive first and second pins 71, 81 to a radial temperature gradient of the working gas that impinges on the aerofoil and thus minimising the temperature gradient in the wall of the aerofoil.

Furthermore, and for the same reasons, the width or the cross-sectional area of the pins 71, 81 may be increased to increase cooling.

Furthermore, the location of each consecutive crossing point P does not need to be the same and can be offset from one crossing point to the next. This would be apparent where the positions of the pins 71, 81 changes along the length of the channel 6. For example, in FIG. 16, which is a view radially inwardly on a cooling channel 6, shows the row of first pins 70 which is twisted. This may be the case where the aerofoil is twisted or arcuate particularly at the leading edge region. The row of second pins 80 may also twist so that consecutive pins 71, 81 are aligned at a preferable angle to one another. As is known in the art the aerofoil is twisted due to preferable aerodynamic reasons. The pins 71 join the aerofoil's wall at the leading edge region and specifically at the aerodynamic leading edge 91; however, where an array of cooling holes exists through the wall to provide a cooling film over the outer surface of the aerofoil, the pins 71 may join the leading edge region away from the aerodynamic leading edge.

While the present technique has been described in detail with reference to certain embodiments, it should be appreciated that the present technique is not limited to those precise embodiments. Rather, in view of the present disclosure which describes exemplary modes for practicing the invention, many modifications and variations would present themselves, to those skilled in the art without departing from the scope and spirit of this invention. The scope of the invention is, therefore, indicated by the following claims rather than by the foregoing description. All changes, modifications, and variations coming within the meaning and range of equivalency of the claims are to be considered within their scope.

Claims

Patent claims

1. A turbomachine component (1) having an aerofoil (5), particularly a blade or a vane for a gas turbine engine (10), the turbomachine component (1) comprising:

- a suction side wall (2) of the aerofoil (5) and a pressure side wall (3) of the aerofoil (5) bordering an aerofoil cavity ( 4 ) , and

- at least one cooling channel (6) extending inside at least a part of the aerofoil cavity (4), wherein the cooling channel (6) is defined by an internal wall (60) of the cooling channel (6) and is adapted to be flowed through by a cooling fluid (7) and wherein the cooling channel (6)

comprises an inlet (66) for receiving the cooling fluid (7) to be flowed through the cooling channel (6),

characterized in that the turbomachine component (1)

comprises

- at least a first row (70) of pins comprising a plurality of first pins (71), wherein one end of each first pin (71) is connected to the internal wall (60) of the cooling channel (6) and another end of each first pin (71) is connected to the internal wall (60) of the cooling channel (6) on opposite side such that each first pin (71) extends within the cooling channel (6); and

- at least a second row (80) of pins comprising a plurality of second pins (81), wherein one end of each second pin (81) is connected to the internal wall (60) of the cooling channel (6) and another end of each second pin (81) is connected to the internal wall (60) of the cooling channel (6) on opposite side such that each second pin (81) extends within the cooling channel (6);

- wherein the first pins (71) and the second pins (81), with respect to each other, are non-contiguously and alternatively arranged within the cooling channel (6) and are disposed at an angle (99); and

- wherein each of the first pins (71) have a centreline () and a length (LI) and each of the second pins (81) have a centreline (CL2) and a length (L2), the centrelines (CL1, CL2) cross each other at a location within at least 10% of one of the lengths (LI, L2) from the internal wall (60) .

2. The turbomachine component (1) according to claim 1, wherein the angle (99) is between 20 degree and 90 degree.

3. The turbomachine component (1) according to claim 1 or 2, wherein the angle (99) is between 45 degree and 90 degree.

4. The turbomachine component (1) according to any of claims 1 to 3, wherein the angle (99) is 90 degree.

5. The turbomachine component (1) according to any of claims 1 to 4, wherein an angle (100) formed by at least one of the first and the second pins (71, 81) with respect to a

direction of flow of the cooling fluid (7) is between 20 degree and 160 degree, and particularly is 90 degree.

6. The turbomachine component (1) according to any of claims 1 to 5, wherein the internal wall (60) of the cooling channel

(6) has a plurality of sections arranged circumferentially with respect to the cooling channel (6), the plurality of sections includes:

- at least one pressure side wall section (61) formed by a part of the pressure side wall (3) ;

- at least one suction side wall section (62) formed by a part of the suction side wall (2); and

- at least one rib section (63) formed by a rib (75) disposed within the aerofoil cavity (4) between the suction side wall (2) and the pressure side wall (3) .

7. The turbomachine component (1) according to claim 6, wherein at least one of the first pins (71) of the first row (70) and the second pins (81) of the second row (80) extend between the pressure side wall section (61) and the suction side wall section (62) of the internal wall (60) of the cooling channel (6) .

8. The turbomachine component (1) according to claim 6 or 7, wherein at least one of the first pins (71) of the first row (70) and the second pins (81) of the second row (80) extend between the rib section (63) and the pressure side or the suction side wall section (61,63) of the internal wall (60) of the cooling channel (6) .

9. The turbomachine component (1) according to claim 6 or 7, wherein the plurality of sections includes two oppositely facing rib sections (63,64) and wherein at least one of the first pins (71) of the first row (70) and the second pins (81) of the second row (80) extend between the two rib sections (63,64) of the internal wall (60) of the cooling channel ( 6) .

10. The turbomachine component (1) according to claim 6, wherein the pressure side wall section (61) and the suction side wall section (62) meet forming an angular section (65) of the internal wall (60) of the cooling channel (6), and wherein one of the first pins (71) of the first row (70) and the second pins (81) of the second row (80) extend between the rib section (63) and the angular section (65) of the internal wall (60) of the cooling channel (6) .

11. The turbomachine component (1) according to claim 10, wherein one of the first pins (71) of the first row (70) and the second pins (81) of the second row (80) extend between the pressure side wall section (61) and the suction side wall section (62) of the internal wall (60) of the cooling channel (6) .

12. The turbomachine component (1) according to claim 10, wherein one of the first pins (71) of the first row (70) and the second pins (81) of the second row (80) extend between the rib section (63) and the pressure side wall section (61) or the suction side wall section (62) of the internal wall (60) of the cooling channel (6) .

13. The turbomachine component (1) according to any of claims 1 to 12, wherein the first pins (71) and/or the second pins (81) are formed as one-part extension of the internal wall (60) of the cooling channel (6) .

14. The turbomachine component (1) according to any of claims 1 to 12, wherein the first pins (71) and/or the second pins (81) are formed as a fixture attached to the internal wall (60) of the cooling channel (6) .

15. The turbomachine component (1) according to any of claims 1 to 14, wherein the alternatively arranged first pins (71) and second pins (81) are equidistantly spaced apart from one another .