EP3460190A1

EP3460190A1 - Heat transfer enhancement structures on in-line ribs of an aerofoil cavity of a gas turbine

Info

Publication number: EP3460190A1
Application number: EP17192319.6A
Authority: EP
Inventors: Anthony Davis
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2017-09-21
Filing date: 2017-09-21
Publication date: 2019-03-27

Abstract

A gas turbine component (1), such as a blade (38) or a vane (40, 44), having an aerofoil (60) is presented. The aerofoil has an interior surface (62) defining an aerofoil cavity (65) configured to be flowed through by cooling air. On the interior surface of the aerofoil, primary ribs extend longitudinally within the airfoil cavity along a flow direction (9) of the cooling air. Each primary rib includes a plurality of secondary heat transfer structures (80) positioned on a surface (72) of the primary ribs. The secondary heat transfer structures may be secondary ribs (82), pin-fins (90), dimples or pimples (95). A surface of the secondary heat transfer structures being continuous with the surface of the primary ribs results in an increase of the surface area of the primary ribs consequently resulting in enhancement in heat transfer, and thereby in resultant cooling, associated with the primary ribs.

Description

The present invention relates to gas turbine engines, and more particularly to gas turbine components having an aerofoil, such as blades or vanes, with in-line ribs.
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 component having an aerofoil, i.e. blade or vane, is to have one or more internal passages, referred to as cooling channels or cooling passages, formed within an aerofoil cavity defined by an interior surface of the aerofoil and to flow a cooling fluid, such as cooling air, through the aerofoil cavity or through the cooling channels formed within the aerofoil cavity thereby cooling the interior surface of the aerofoil, and consequently cooling the aerofoil of the blade/vane.
Generally, the cooling fluid enters the aerofoil, i.e. the cooling air flows into the aerofoil cavity, though an opening of the aerofoil cavity from one radial side of the aerofoil such as a platform side or a root side in case of blades or vanes, or from both sides such as a tip and a root side in case of vanes, and then flows along and in contact with the interior surface, often in a serpentine flow path defined within the aerofoil cavity, and exits the aerofoil cavity for film cooling through holes that open on an external or exterior surface of the aerofoil or exits through slots formed, generally on a trailing edge of the aerofoil. Alternately the cooling air is flowed through an impingement tube positioned within the aerofoil cavity and exited out from the impingement tube via impingement holes of the impingement tube in form of impingement jets directed towards the interior surface of the aerofoil. The cooling air after impinging on the interior surface of the aerofoil flows along and in contact with the interior surface, and exits the aerofoil cavity for film cooling through holes that open on the external surface of the aerofoil or exits through slots formed, generally at or near the trailing edge of the aerofoil.
To direct the flow of the cooling air along the interior surface of the aerofoil, ribs are formed on the interior surface, and agreeing with the direction of flow of the cooling air extend chordwise or spanwise within the interior surface. The chordwise extending ribs extend from a leading edge side towards the trailing edge side of the aerofoil and direct the cooling air towards the trailing edge, when the impingement jets impact the interior surface at the leading edge side. The chordwise extending ribs also assist in locating the impingement tube within the aerofoil cavity. The spanwise extending ribs are often formed within the cooling channels and direct the cooling air radially inwards or radially outwards, with respect to a rotational axis of the gas turbine, within the aerofoil cavity i.e. from the tip end of the aerofoil towards the platform end of the aerofoil and vice versa. It may be noted that in both aforementioned schemes of ribs arrangement the ribs extend longitudinally along the direction of flow of the cooling air, and hence the aforementioned spanwise or chordwise extending ribs are generally referred to as in-line ribs signifying their arrangement with respect to the direction of flow of the cooling air i.e. being in-line with the direction of flow of the cooling air. Thus in other words, the aforementioned ribs, whether extending chordwise or extending spanwise, are aligned longitudinally on the interior surface of the aerofoil along the intended flow direction of the cooling air. The ribs besides directing the cooling air increase a surface area of the interior surface and thus result in increased heat transfer from the interior surface to the cooling air flowing over, i.e. flowing along and in contact with, the interior surface. However, further enhancement of heat transfer and thus an increase in cooling of the aerofoil is advantageous. Therefore, a technique is desired to increase heat transfer from interior surface of the aerofoil and thus increase the cooling of the aerofoil.
Thus the object of the present invention is to provide a technique for increasing heat transfer from interior surface of the aerofoil to the cooling air flowing over the interior surface and thus increasing cooling of the aerofoil.
The above objects are achieved by a gas turbine component according to claim 1 of the present technique. Advantageous embodiments of the present technique are provided in dependent claims.
The invention presents a gas turbine component having enhanced cooling features on a surface of the in-line ribs. The gas turbine component, hereinafter also referred to as the component, has an aerofoil having an aerofoil cavity defined by an interior surface of the aerofoil. The component may be a blade or a vane of a gas turbine. The aerofoil cavity is configured to be flowed through by cooling air. The component includes a plurality of primary ribs arranged on the interior surface of the aerofoil and within the aerofoil cavity. The primary ribs are spaced apart and extend from the interior surface towards an inside of the aerofoil cavity.
The primary ribs extend longitudinally along a flow direction of the cooling air on the interior surface of the aerofoil, i.e. the primary ribs are the in-line ribs as explained hereinabove. In other words, the primary ribs arise from the interior surface of the aerofoil and have a free-hanging end disposed in the aerofoil cavity. In the component, each of the primary ribs comprises a plurality of secondary heat transfer structures. The secondary heat transfer structures are positioned on a surface of the primary ribs i.e. the secondary heat transfer structures, hereinafter also referred to as the secondary structures, extend from the surface of the primary ribs and have a free-hanging end disposed in the aerofoil cavity. A surface of the secondary structures being continuous with the surface of the primary ribs results in an increase of the surface area of the primary ribs consequently resulting in increase in heat transfer, and thereby in resultant cooling, associated with the primary ribs. The secondary structures are the enhanced cooling features.
The phrase 'longitudinally' with reference to the primary ribs as used herein means along a longitudinal axis of the primary ribs i.e. along a long axis or a lengthwise disposed axis of the primary rib. The longitudinal axis of the primary rib is generally parallel to the interior surface of the aerofoil and may follow a contour of the interior surface. The phrase 'extend longitudinally along a flow direction of the cooling air' means the longitudinal axis of the primary rib is aligned parallel to the flow direction of the cooling air.
In an embodiment of the component, the primary ribs extend longitudinally between a tip end and a platform end of the aerofoil i.e. the longitudinal axis of the primary ribs extends between the tip end and the platform end i.e. extends from the tip end to the platform end. In other words, the primary ribs extend radially, with respect to a rotation axis of the gas turbine. The cooling air entering the aerofoil cavity, hereinafter also referred to as the cavity, generally enters through the platform end of the aerofoil or in other words from a root section in case where the component is a blade of the gas turbine engine and flows within the aerofoil cavity towards the tip end of the aerofoil. In case where the component is a vane of a gas turbine engine, the cooling air enters from a root section of the vane and/or a tip section of the vane and flows within the aerofoil of the vane generally radially inwards or outwards, respectively. The cooling air generally flows in a cooling channel having a serpentine path formed by one or more hairpin turns of the cooling channel. The primary ribs extend longitudinally between one or more sections of the cooling channel between the hairpin turns.
In an embodiment of the component, the primary ribs extend longitudinally between a leading edge and a trailing edge of the aerofoil i.e. the longitudinal axis of the primary ribs extends between the leading edge and the trailing edge i.e. extends from a side of the leading edge to a side of the trailing edge. Such ribs are generally used to locate an impingement tube within the aerofoil cavity. Impingement jets ejected from the impingement tube, particularly from impingement holes present in the impingement tube, are directed towards the interior wall of the cavity. However after encountering the interior wall of the cavity the cooling air is directed generally in a flow direction from the leading edge towards the trailing edge by the primary ribs that extend longitudinally between the leading edge and the trailing edge. The flow may also be directed in a general direction towards the leading edge if appropriate.The cooling air after flowing towards the trailing edge is ejected out of the aerofoil through holes or slots for example holes or slots at the trailing edge, holes or slots at the leading edge. Besides increasing the surface area for heat exchange between the cooling air and the primary ribs, the secondary ribs also increase the turbulence in the cooling air.
In another embodiment of the gas turbine component, the secondary heat transfer structures are secondary ribs. Ribs are devices or features that emulate a rod or a bar i.e. a rigid strip shaped and having a cross-section that is semicircular, parabolic or polygonal.
In another embodiment of the gas turbine component, the secondary ribs extend longitudinally along the flow direction of the cooling air i.e. a longitudinal axis of the secondary rib is parallel to the longitudinal axis of the primary rib. The phrase 'longitudinally' with reference to the secondary ribs as used herein means along the longitudinal axis of the secondary ribs i.e. along a long axis or a lengthwise disposed axis of the secondary rib. The longitudinal axis of the secondary rib for this embodiment is generally parallel to the interior surface of the aerofoil.
In another embodiment of the gas turbine component, the secondary ribs extend longitudinally at an angle to the flow direction of the cooling air i.e. the longitudinal axis of the secondary rib is at an angle to the longitudinal axis of the primary rib. The phrase 'longitudinally' with reference to the secondary ribs as used herein means along the longitudinal axis of the secondary ribs i.e. along a long axis or a lengthwise disposed axis of the secondary rib. The longitudinal axis of the secondary rib for this embodiment is at the aforementioned angle to the interior surface of the aerofoil. In a related embodiment the angle is a right angle, or in other words the longitudinal axis of the secondary ribs is mutually perpendicular to the longitudinal axis of the primary rib on which the secondary ribs are positioned and to the interior surface of the aerofoil.
In another embodiment of the gas turbine component, a maximum cross-sectional extent of the secondary ribs is between 50% and 100% of a thickness of the primary rib at which the secondary ribs are positioned, particularly about 75% of the thickness of the primary rib at which the secondary ribs are positioned. A height of the secondary ribs is between one time and two and half times of the maximum cross-sectional extent of the secondary rib.
In another embodiment of the gas turbine component, the secondary heat transfer structures are pin-fins. In a related embodiment, a diameter of the pin-fins is between 50% and 100% of a thickness of the primary rib at which the pin-fins are positioned, particularly about 75% of the thickness of the primary rib at which the pin-fins are positioned, and a height of the pin-fins is between one time and two and half times of the diameter of the pin-fin.
In another embodiment of the gas turbine component, the secondary heat transfer structures are pimples or dimples. In a related embodiment, a maximum cross-sectional extent of the pimples or dimples is between 50% and 100% of a thickness of the primary rib at which the pimples or dimples are positioned, and particularly about 75% of the thickness of the primary rib at which the pimples or dimples are positioned.
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 gas turbine engine in a sectional view in which a gas turbine component of the present technique is incorporated;
FIG 2: schematically illustrates an exemplary embodiment of the gas turbine component of the present technique, particularly depicting an aerofoil of the gas turbine component of the present technique with chordwise extending in-line ribs i.e. an exemplary embodiment of primary ribs of the present technique;
FIG 3: schematically illustrates a detailed view of a region of the aerofoil of FIG 2 without depicting secondary heat transfer structures;
FIG 4: schematically illustrates conventionally known inline ribs extending chordwise;
FIG 5: schematically illustrates, in contrast to FIG 4, an exemplary embodiment of primary ribs, i.e. in-line ribs, of the present technique depicting secondary heat transfer structure, embodied as secondary ribs, disposed on a surface of the primary ribs in accordance with the present technique;
FIG 6: schematically illustrates another exemplary embodiment of the gas turbine component of the present technique, particularly depicting the aerofoil of the gas turbine component of the present technique with spanwise extending in-line ribs i.e. another exemplary embodiment of primary ribs of the present technique;
FIG 7: schematically illustrates conventionally known inline ribs extending spanwise;
FIG 8: schematically illustrates, in contrast to FIG 7, an exemplary embodiment of the primary ribs, i.e. inline ribs, of the present technique depicting the secondary heat transfer structure, embodied as the secondary ribs, disposed on the surface of the primary ribs in accordance with the present technique;
FIG 9: schematically represents an exemplary embodiment of a flow of cooling air through in-between two primary ribs of the present technique;
FIG 10: schematically represents another exemplary embodiment of the flow of cooling air through inbetween two primary ribs of the present technique;
FIG 11: schematically represents exemplary embodiments of shapes and arrangement of the secondary ribs on the surface of the primary ribs in accordance with the present technique;
FIG 12: schematically represents further exemplary embodiments of shapes and arrangement of the secondary ribs on the surface of the primary ribs in accordance with the present technique;
FIG 13: schematically represents another exemplary embodiment of an arrangement of the secondary ribs on the surface of the primary ribs in accordance with the present technique;
FIG 14: schematically represents another exemplary embodiment of the secondary heat transfer structures embodied as pin-fins in accordance with the present technique; and
FIG 15: schematically represents yet another exemplary embodiment of the secondary heat transfer structures embodied as pimples or dimples in accordance with the present technique.

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 exemplary embodiment of a gas turbine engine 10 in a sectional view in which a gas turbine component 1 (shown in FIGs 2,5,6,8 and 9 to 15) is incorporated. The gas turbine engine 10, hereinafter referred to as the 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 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 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, 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. An inner surface 55 of the transition duct 17 defines a part of the hot gas path.
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 section 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 are depicted, 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 that turn the flow of working gas onto the turbine blades 38.
The combustion gas 34 from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22. The guiding vanes 40, 44, hereinafter also referred to as the vanes 40,44, serve to optimise the angle of the combustion or working gas 34 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 operations 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. Furthermore, the cannular combustor section arrangement 16 is also used for exemplary purposes and it should be appreciated that the present technique is equally applicable to annular type and can type combustion chambers.
The terms axial, axially, axial direction, radial, radially, radial direction, circumferential, circumferentially and circumferential direction, etc are made with reference to the rotational axis 20 of the gas turbine engine 10, unless otherwise stated.
The gas turbine component 1 of the present technique may be the blade 38 as shown in FIG 1 or the vane 40,44 as shown in FIG 1. It may be noted that the present technique has been primarily explained in details hereinafter with respect to an embodiment of the turbine blade 38 as the gas turbine component 1 of the present technique, however, it must be appreciated that the present technique is equally applicable and implemented similarly with respect to the turbine vane 40,44 or any other turbomachine component having an aerofoil and being cooled by a cooling fluid flowing through an aerofoil cavity formed within the aerofoil.
FIGs 2, 3 and 5 are used hereinafter to describe an exemplary embodiment of a gas turbine engine component 1 of the present technique. FIG 4 has been used to depict a conventional design and for explaining difference between the conventional design as shown in FIG 4 and the gas turbine engine component 1 of the present technique as shown in FIG 5.
The gas turbine component 1, hereinafter also referred to as the component 1, has an aerofoil 60. The aerofoil 60 may extend from a platform (not shown in FIG 2) in the radial direction. From the platform also emanates a root (not shown in FIG 2) or a fixing part generally when the component 1 is embodied as the blade 38 (shown in FIG 1). The root or the fixing part may be used to attach the component 1 to the turbine disc 38 (shown in FIG 1). It may be noted that in some other embodiments of the component 1, for example when the component 1 is the vane 40,44 the aerofoil 60 may be affixed to fixed structures such as the stator discs or the housing 50, as shown in FIG 1.
The aerofoil 60 includes a pressure side 67 and a suction side 68. The pressure side 67 and the suction side 68 meet at a trailing edge 66 on one end and a leading edge 64 on another end. The aerofoil 60 has a platform end 63 at which the aerofoil 60 is attached to the platform, and a tip end 61, radially outwards of the platform end 63. The aerofoil 60 may be connected to a shroud (not shown in FIG 2) at the tip end 61 of the aerofoil 60. In some other embodiments the aerofoil 60 may be connected to a tip platform (not shown) instead of the shroud. The pressure side 67, the suction side 68, the leading edge 64 and the trailing edge 66 together have a continuously extending interior surface 62 and a continuously extending exterior surface 69 or an external surface 69. The interior surface 62 of the aerofoil 60 binds or defines an aerofoil cavity 65 formed within the aerofoil 60. The aerofoil cavity 65, hereinafter also referred to as the cavity 65, extends radially or spanwise between the tip end 61 and the platform end 63 of the aerofoil 60.
Within the cavity 65 an impingement tube 2 may be located. A region of the aerofoil 60 marked with reference character A in FIG 2 has been shown in details in FIG 3. To cool the interior surface 62 of the aerofoil 60, cooling air is made to flow through the impingement tube 2 located within the cavity 65 and in vicinity of the interior surface 62. The cooling air exits the impingement tube 2 via impingement holes 3 of the impingement tube 2 in form of impingement jets 4 directed towards the interior surface 62 of the aerofoil 60. After impacting the interior surface 62 the cooling air is guided by ribs 70, referred to as primary ribs 60, formed on the interior surface 62 of the aerofoil 60. The cooling air is guided along the interior surface 62 in a direction extending from the leading edge 64 towards the trailing edge 66. The direction of flow of cooling air has been depicted in FIGs 2 and 3 by arrows marked with reference numeral 9. In another embodiment (not shown), instead of or in addition to being directed towards the trailing edge 66, the cooling air is directed towards the leading edge 64.
In the component 1, a plurality of the primary ribs 70 are arranged on the interior surface 62 of the aerofoil 60. The primary ribs 70 are spaced apart, generally radially, from each other and extend from the interior surface 62 towards an inside of the aerofoil cavity 65. The primary ribs 70 extend longitudinally along the flow direction 9 of the cooling air. In the example of FIGs 2, 3 and 5 the primary ribs 70 thus extend in the direction extending from the leading edge 64 towards the trailing edge 66.
The phrase 'longitudinally' with reference to the primary ribs 70 as used herein means along a longitudinal axis 71 of the primary ribs 70 i.e. along a long axis 71 or a lengthwise disposed axis 71 of the primary rib 70. The longitudinal axis 71 of each of the primary ribs 70 is generally parallel to the interior surface 62 of the aerofoil 60 and may follow a contour of the interior surface 62. In other words, the longitudinal axis 71 of each of the primary ribs 70 is parallelly aligned with the flow direction 9 of the cooling air as shown in FIG 3. The primary ribs 70 are thus chordwise extending i.e. extending longitudinally between the leading edge 64 and the trailing edge 66 of the aerofoil 60, or in other words the longitudinal axis 71 of the primary ribs 70 extends chordwise or between the leading edge 64 and the trailing edge 66 of the aerofoil 60.
In the component 1 as shown in FIG 5, each of the primary ribs 70 comprises a plurality of secondary heat transfer structures 80. The secondary heat transfer structures 80 are positioned on a surface 72 of the primary ribs 70 i.e. the secondary heat transfer structures 80, hereinafter also referred to as the secondary structures 80, extend from the surface 72 of the primary ribs 70 and have a free-hanging end disposed in the aerofoil cavity 65. The secondary structures 80 are devices or features or elements that result in increase in a surface area of the surface 72 of the primary ribs 70. The secondary structures 80 may be rib shaped, and therefore referred to as secondary ribs 82 as shown in FIG 5, or may have other shapes as explained later with reference to FIGs 14 and 15.
FIG 4 presents conventionally known chordwise extending ribs 70. However unlike the example of FIG 5, the conventionally known chordwise extending ribs 70 do not have the secondary structures 80 on their surface, and thus the surface 72 of the conventionally known chordwise extending ribs 70 is flat or smooth as shown in FIG 4. Whereas in the component 1 of the present technique as shown in FIG 5, the surface 72 of the primary ribs 70, also extending chordwise, has the secondary structures 80 formed on the surface 72 of the primary ribs 70.
FIGs 6 and 8 are used hereinafter to describe another exemplary embodiment of the gas turbine engine component 1 of the present technique. FIG 7 has been used to depict a conventional design and for explaining difference between the conventional design as shown in FIG 7 and the gas turbine engine component 1 of the present technique as shown in FIG 8.
The component 1 has the aerofoil 60 extending from a platform 7 in the radial direction. The platform end 63 of the aerofoil 60 is the part of the aerofoil 60 at which the aerofoil 60 is attached to the platform 7. The tip end 61 is the radially outward end of the aerofoil 60, and opposite to the platform end 63. From the platform 7 also emanates the root 8 generally when the component 1 is embodied as the blade 38 (shown in FIG 1). Within the aerofoil 60 the interior surface 62 defines the aerofoil cavity 65. The aerofoil cavity 65 may be designed such that, by forming cooling channels within the cavity 65, the cooling air enters the aerofoil cavity 65 from the platform end 63 and flows within the cavity 65 in a serpentine path as depicted by the arrows marked with reference numeral 9 in FIG 6.
The cooling air after entering the cavity 65 from the platform end 63 flows radially outwardly i.e. towards the tip end 61 in a part of the cooling channel formed by dividing the cavity 65 spanwise. The cooling channel thereafter makes a hairpin turn and thus the cooling air flows within the cavity 65 from the tip end 61 towards the platform end 63. The cooling channel may make further hairpin turns thus making the cooling air flow in the serpentine path. It may be noted that the cooling channel depicted in FIG 6 is for exemplary purposes only. It may be appreciated by one skilled in the art that the cooling channel may have more tortuous path than the one depicted in FIG 6. It may also be appreciated by one skilled in the art that the cooling channel may have less tortuous path than the one depicted in FIG 6, for example the cooling path may be the cavity 65 without any hairpin bends.
A region of the aerofoil 60 marked with reference character B in FIG 6 has been shown in details in FIG 8. In the component 1 shown in example of FIG 6, the primary ribs 70 are arranged on the interior surface 62 of the aerofoil 60. The primary ribs 70 are spaced apart, chordwise, from each other and extend from the interior surface 62 towards the inside of the aerofoil cavity 65. The primary ribs 70 extend longitudinally along the flow direction 9 of the cooling air. In the example of FIGs 6 and 8 the primary ribs 70 thus extend in the direction extending between the tip end 61 and the platform end 63 of the aerofoil 60.
The longitudinal axis 71 of the primary ribs 70 is aligned extending between the tip end 61 and the platform end 63 of the aerofoil 60 and is generally parallel to the interior surface 62 of the aerofoil 60, and may follow a contour of the interior surface 62. In other words, the longitudinal axis 71 of each of the primary ribs 70 is parallelly aligned with the flow direction 9 of the cooling air as shown in FIGs 6 and 8. The primary ribs 70 are thus spanwise extending i.e. extending longitudinally between the tip end 61 and the platform end 63 of the aerofoil 60, or in other words the longitudinal axis 71 of the primary ribs 70 extends spanwise.
In the component 1 as shown in FIG 8, each of the primary ribs 70 comprises the secondary structures 80. The secondary structures 80 are positioned on the surface 72 of the primary ribs 70 i.e. the secondary structures extend from the surface 72 of the primary ribs 70 and have a free-hanging end disposed in the aerofoil cavity 65. The secondary structures 80 are devices or features or elements that result in increase in a surface area of the surface 72 of the primary ribs 70. The secondary structures 80 may be rib shaped, and therefore referred to as the secondary ribs 82 as shown in FIG 8, or may have other shapes as explained later with reference to FIGs 14 and 15.
FIG 7 presents conventionally known spanwise extending ribs 70. However unlike the example of FIG 8, the conventionally known spanwise extending ribs 70 do not have the secondary structures 80 on their surface, and thus the surface 72 of the conventionally known spanwise extending ribs 70 is flat or smooth as shown in FIG 7. Whereas in the component 1 of the present technique as shown in FIG 8, the surface 72 of the primary ribs 70, also extending spanwise, has the secondary structures 80 formed on the surface 72 of the primary ribs 70.
The secondary structures 80 help in increasing the surface area for heat transfer between the primary ribs 70 and the cooling air. Additionally, the secondary structures 80 also help in inducing turbulence in the cooling air flow 9. FIGs 9 and 10 show the cooling air flowing between two of the primary ribs 70. FIG 9 shows the secondary structures 80 having rectangular cross-sections as for the secondary ribs 82 of FIGs 5 and 8. FIG 10 shows the secondary structures 80 having semi-circular cross-sections instead of the rectangular cross-section shown in FIGs 5 and 8. The cooling air flow 9 enters non-turbulently, i.e. emulating laminar flow, but due to the secondary structures 80 present on the surface 72 of the primary ribs 70 turbulence 5 is generated in the cooling air.
Hereinafter different shapes of the secondary structures 80 when embodied as the secondary ribs 82 and their relative orientation or arrangement with respect to the primary ribs 70 and/or the interior surface 62 and/or the flow direction 9 of the cooling air have been explained with reference to FIGs 11, 12 and 13.
As shown in FIGs 11 and 12, in an embodiment of the gas turbine component 1, the secondary ribs 82 are formed on the surface 72 of the primary ribs 70. The secondary ribs 82 are devices or features that emulate a rod or a bar i.e. a rigid strip shaped and having a cross-section that is semicircular, parabolic or polygonal. FIG 12 shows three exemplary shapes of the secondary ribs 82 - a secondary rib 82a having a rectangular cross-section, another secondary rib 82b having a semi-circular or parabolic cross-section, and yet another secondary rib 82c having a triangular cross-section. It may be noted that further shapes of the secondary ribs 82 are also within the scope of the present technique, for example a pentagonal or hexagonal cross-section of the secondary ribs 82.
In an embodiment of the gas turbine component 1 as depicted in FIG 13, the secondary ribs 82 extend longitudinally along the flow direction 9 of the cooling air i.e. a longitudinal axis 81 of the secondary rib 82 is parallel to the longitudinal axis 71 of the primary rib 70. The phrase 'longitudinally' with reference to the secondary ribs 82 as used herein means along the longitudinal axis 81 of the secondary ribs 82 i.e. along a long axis 81 or a lengthwise disposed axis 81 of the secondary rib 82. The longitudinal axis 81 of the secondary rib 82 for this embodiment is generally parallel to the interior surface 62 of the aerofoil 60.
In another embodiment of the gas turbine component 1 as shown in FIGs 11 and 12, the secondary ribs 82 extend longitudinally at an angle 75 to the flow direction 9 of the cooling air i.e. the longitudinal axis 81 of the secondary rib 82 is at the angle 75 to the longitudinal axis 71 of the primary rib 70. The phrase 'longitudinally' with reference to the secondary ribs 82 as used herein means along the longitudinal axis 81 of the secondary rib 82 i.e. along the long axis 81 or the lengthwise disposed axis 81 of the secondary rib 82. The longitudinal axis 81 of the secondary rib 82 for this embodiment is at the aforementioned angle 75 to the interior surface 62 of the aerofoil 60. The angle 75 may be 10 degree to 90 degree. In a related embodiment as shown in FIG 12, the angle 75 is a right angle, or in other words the longitudinal axis 81 of each of the secondary ribs 82 is mutually perpendicular to the longitudinal axis 71 of the primary rib 70 on which the secondary ribs 82 are positioned and to the interior surface 62 of the aerofoil 60.
It may be noted that the heat transfer to the cooling air from the secondary rib 82 varies with change in the angle 75, i.e. for example measure of heat transfer expressed via Nusselt number is smaller when the secondary ribs 82 are at right angle to the flow direction 9 of the cooling air compared to the heat transfer when the secondary ribs 82 are at another angle, for example 45 degree to the flow direction 9 of the cooling air. The angle 75 is the smaller angle of the two complementary angles formed by the secondary rib 82 with the flow direction 9 of the cooling air, or in other words with the longitudinal axis 71 of the primary rib 70. However, the increase in heat transfer also results in increase in pressure loss in the cooling air. So depending on the level of heat transfer required, and the pressure available, the choice of the angle 75 for orienting the secondary ribs 82 on the surface 72 of the primary rib 70 can be made.
As shown in FIG 11, a maximum cross-sectional extent L of the secondary ribs 82 is between 50% and 100% of a thickness T of the primary rib 70 at which the secondary ribs 82 are positioned, particularly about 75% of the thickness T of the primary rib 70 at which the secondary ribs 82 are positioned. The maximum cross-sectional extent L for a given secondary rib 82 means greatest or maximum measure of length, for the given secondary rib 82, performed parallel to the surface 72 of the primary rib 70 and perpendicular to the longitudinal axis 81 of the secondary rib 82. In other words, the maximum cross-sectional extent L for a given secondary rib 82 means a measure of length, for the given secondary rib 82, performed parallel to the surface 72 of the primary rib 70 and perpendicular to the longitudinal axis 81 of the secondary rib 82. The thickness T of the primary rib is measured perpendicular to the surface 72 of the primary rib 70. A height H of each of the secondary ribs 82 is between one time, i.e. equal to, and two and half times (2.5 times) of the maximum cross-sectional extent L of the secondary rib 82. The height H is measured perpendicular to the surface 72 of the primary rib 70.
Hereinafter referring to FIGs 14 and 15, embodiments of the secondary structure 80, other than the secondary ribs 82, have been explained. In an embodiment of the component 1 as schematically depicted in FIG 14, the secondary structures 80 are pin-fins 90. As shown in FIG 14, a diameter d of the pin-fins 90 is between 50% and 100% of the thickness T, as aforementioned, of the primary rib 70 at which the pin-fins 90 are positioned, and particularly about 75% of the thickness T of the primary rib 70 at which the pin-fins 90 are positioned. A height h of the pin-fins 90 is between one time, i.e. equal to, and two and half times (2.5 times) of the diameter d of the pin-fin 90.
In another embodiment of the component 1 as schematically depicted in FIG 15, the secondary structures 80 are pimples 95 or dimples 95. A maximum cross-sectional extent L' of the pimples 95 or the dimples 95 is between 50% and 100% of the thickness T of the primary rib 70 at which the pimples 95 or dimples 95 are positioned, and particularly about 75% of the thickness T of the primary rib 70 at which the pimples 95 or the dimples 95 are positioned. The maximum cross-sectional extent L' of the pimples 95 or the dimples 95 is the measure of a diameter of the pimple 95 or the dimple 95 performed at a level of the surface 72 of the primary rib 70.
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

A gas turbine component (1) comprising:
- an aerofoil (60) having an aerofoil cavity (65) defined by an interior surface (62) of the aerofoil (60), wherein the aerofoil cavity (65) is configured to be flowed through by cooling air;

- a plurality of primary ribs (70) arranged spaced apart on the interior surface (62) of the aerofoil (60) and within the aerofoil cavity (65), wherein the primary ribs (70) extend longitudinally on the interior surface (62) of the aerofoil (60) along a flow direction (9) of the cooling air,
characterized in that
each of the primary ribs (70) comprises a plurality of secondary heat transfer structures (80), wherein the secondary heat transfer structures (80) are positioned on a surface (72) of the primary ribs (70).
The gas turbine component (1) according to claim 1, wherein the primary ribs (70) extend longitudinally between a tip end (61) and a platform end (63) of the aerofoil (60).
The gas turbine component (1) according to claim 1, wherein the primary ribs (70) extend longitudinally between a leading edge (64) and a trailing edge (66) of the aerofoil (60).
The gas turbine component (1) according to any of claims 1 to 3, wherein the secondary heat transfer structures (80) are secondary ribs (82).
The gas turbine component (1) according to claim 4, wherein the secondary ribs (82) extend longitudinally along the flow direction (9) of the cooling air.
The gas turbine component (1) according to claim 4, wherein the secondary ribs (82) extend longitudinally at an angle (75) to the flow direction (9) of the cooling air.
The gas turbine component (1) according to claim 5, wherein the angle (75) is a right angle.
The gas turbine component (1) according to any of claims 4 to 7, wherein the secondary ribs (82) have one of a semicircular cross-section, a parabolic cross-section, and a polygonal cross-section.
The gas turbine component (1) according to any of claims 4 to 8, wherein a maximum cross-sectional extent (L) of the secondary ribs (82) is between 50% and 100% of a thickness (T) of the primary rib (70) at which the secondary ribs (82) are positioned, particularly about 75% of the thickness (T) of the primary rib (70) at which the secondary ribs (82) are positioned.
The gas turbine component (1) according to any of claims 1 to 3, wherein the secondary heat transfer structures (80) are pin-fins (90).
The gas turbine component (1) according to claim 10, wherein a diameter (D) of the pin-fins (90) is between 50% and 100% of a thickness (T) of the primary rib (70) at which the pin-fins (90) are positioned, particularly about 75% of the thickness (T) of the primary rib (70) at which the pin-fins (90) are positioned, and wherein a height (H) of the pin-fins (90) is between one time and two and half times of the diameter (D) of the pin-fin (90).
The gas turbine component (1) according to any of claims 1 to 3, wherein the secondary heat transfer structures (80) are pimples or dimples (95).
The gas turbine component (1) according to claim 12, wherein a maximum cross-sectional extent (L') of the pimples or dimples (95) is between 50% and 100% of a thickness (T) of the primary rib (70) at which the pimples or dimples (95) are positioned, particularly about 75% of the thickness (T) of the primary rib (70) at which the pimples or dimples (95) are positioned.
The gas turbine component (1) according to any of claims 1 to 13, wherein the gas turbine component (1) is a blade (38) of a gas turbine engine (10).
The gas turbine component (1) according to any of claims 1 to 13, wherein the gas turbine component (1) is a vane (40,44) of a gas turbine engine (10).