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(11) | EP 3 241 991 A1 |
| (12) | EUROPEAN PATENT APPLICATION |
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| (54) | TURBINE ASSEMBLY |
| (57) A turbine assembly (10, 10a) comprises a hollow aerofoil (12) having at least a main
cavity (14) and a leading edge cavity (16, 16a), at least one shared central wall
(22) connecting the main cavity and the leading edge cavity and comprising at least
one main impingement hole (24), a left side cavity (28), and a right side cavity (30),
wherein each of the left and right side cavities is connected to the leading edge
cavity by a left, respectively a right partitioning wall (32,34), both respectively
positioned at opposed sides of the leading edge cavity in respect to a fictitious
connection (36) of a midpoint (38) of the at least one main impingement hole and the
leading edge and wherein each of the at least one partitioning walls comprises at
least one side impingement hole (40, 42).
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Field of the Invention
[0001] The present invention relates to a turbine assembly such as turbine rotor blades and stator vanes and to impingement structures used in such components for cooling purposes.
Background to the Invention
[0002] The performance of a turbine or a specifically of a gas turbine improves mainly by increasing the gas turbine entry temperature (TET). Hence, modern turbines often operate at extremely high temperatures. The effect of temperature on the turbine blades, stator vanes and surrounding components can be detrimental to the efficient operation of the turbine and can, in extreme circumstances, lead to distortion and possible failure of such components. In order to overcome this risk, cooling systems, made up of a series of cavities in the blades and vanes, may be employed. They are designed to reduce the metal temperature of these components. Air, coming from a compressor at a lower temperature than the one present in a turbine annulus, passes into these ducts and extracts heat flux from the metal and is spent to aid the cooling at a trailing edge or in small holes at a leading edge, pressure side, or suction side, so-called film cooling holes.
[0003] Furthermore, an effective method to cool the leading edge may be the use of impingement cooling in order to have high heat flux. Here the coolant is directed through small holes (in a tube or casted in aerofoil walls), which accelerate the flow and direct it on a surface to be cooled. After hitting the surface and extracting the heat flux thereby the flow is spent in film cooling holes at the leading edge. However, in case film cooling hole in the vane or blade are disadvantageous, because of the technologic class of the blade or vane, the impingement cooling is less effective due to e.g. created transverse flows along the wall to be cooled.
[0004] It is a first objective of the present invention to provide an advantageous aerofoil-shaped turbine assembly such as a turbine rotor blade and a stator vane with which the above-mentioned shortcomings can be mitigated, and especially a more aerodynamic efficient aerofoil and gas turbine component is facilitated.
Summary of the Invention
[0005] Accordingly, the present invention provides a turbine assembly comprising a basically hollow aerofoil having at least a main cavity and a leading edge cavity positioned at a leading edge of the hollow aerofoil, wherein the main cavity and the leading edge cavity extent at least partially in span-wise direction of the hollow aerofoil and wherein the hollow aerofoil comprises at least one shared central wall connecting the main cavity and the leading edge cavity, wherein the at least one central wall comprises at least one main impingement hole, and wherein the leading edge cavity is closed off in respect to an exterior of the hollow aerofoil.
[0006] It is provided that the hollow aerofoil comprises at least a left side cavity and a right side cavity, wherein each side cavity extent at least partially in span-wise direction of the hollow aerofoil and wherein the left side cavity is connected to the leading edge cavity by at least one shared left partitioning wall and the right side cavity is connected to the leading edge cavity by at least one shared right partitioning wall, wherein the at least one left partitioning wall and the at least one right partitioning wall are positioned at opposed sides of the leading edge cavity in respect to a hypothetical or fictitious connection of a midpoint of the at least one main impingement hole and the leading edge and wherein each of the at least one partitioning walls comprise at least one side impingement hole.
[0007] Due to the inventive matter an effective cooling of the leading edge can be provided. Its use can improve the life of the turbine assembly or aerofoil, like a blade or a vane, in absence of film cooling holes by increasing the cooling efficiency. Further, transverse flows that may occur in the leading edge cavity and are usually drained via film cooling holes to reduce the transverse flows can be supressed or at least their negative effects may be mitigated. Additionally, this construction increases the cooling efficiency for a blade and vane without film cooling because the extensive use of impingement decreases the metal temperature in comparison with aerofoils cooled with other convective cooling, like vortex cooling or turbulated channels. In some gas turbine blades and vanes the pressure margin necessary to guarantee the integrity of the component without ingestion can't be achieved by using film cooling holes. Hence, by omitting state of the art film cooling holes at the leading edge of the aerofoil in the inventive design a beneficial performance of the gas turbine engine can be provided. Moreover, a freedom of design for possibly used blades and vanes can be expanded. Furthermore, in the absence of film cooling the gas turbine efficiency can be increased since cooling medium discharged by film cooling is prevented from mixing with the main flow. In addition, a blockage of the film cooling holes due to the environmental dust or the typology of fuel can be prevented. By avoiding the use of film cooling a wider range of fuel can be granted and the turbine can work in very harsh environment. Further, a drop in pressures due to the jet impinging holes in series to connect each cavity is compensated by the pressure ratio between the root of the blade and the trailing edge.
[0008] Even if a term like aerofoil, cavity, wall, hole, surface, channel, aperture, sub-cavity, turbulating structure, dimple, turbulator, fin or platform is used in the singular or in a specific numeral form in the claims and the specification the scope of the patent (application) should not be restricted to the singular or the specific numeral form. It should also lie in the scope of the invention to have more than one or a plurality of the above mentioned structure(s).
[0009] A turbine assembly is intended to mean an assembly provided for a turbine engine, like a gas turbine, wherein the assembly possesses at least an aerofoil. Preferably, the turbine assembly has a turbine wheel or a turbine cascade with circumferential arranged aerofoils. In case of an aerofoil (turbine blade) of a turbine wheel the aerofoil may have a tip and a fire tree root portion and several aerofoils are connected with one another by a disc. In case of a turbine cascade the aerofoil (turbine vane) may have an outer and an inner platform arranged at opponent ends of the aerofoil(s). A span-wise direction of the aerofoil is defined as a direction extending basically perpendicular, preferably perpendicular, to a direction from the leading edge to the trailing edge of the aerofoil.
[0010] In this context a "basically hollow aerofoil" means an aerofoil with a casing, wherein the casing encases at least one cavity. A structure, like a rip, which divides different cavities in the aerofoil from one another and for example extends in a span-wise direction of the aerofoil, does not hinder the definition of "a basically hollow aerofoil". Preferably, the aerofoil is hollow. In particular, the basically hollow aerofoil, referred as aerofoil in the following description, has two cooling regions, an impingement cooling region at a leading edge of the aerofoil and a state of the art pin-fin/pedestal cooling region at the trailing edge. These regions could be separated from one another through a rip.
[0011] In this context the term "shared" should be understood in that the respective cavities comprise or are restricted at least partially by the same wall. Moreover, the terms "central" and "partitioning wall" should be understood that the respective wall is a wall positioned inside the volume of the aerofoil and is no outer wall of the aerofoil that is in direct contact with an exterior of the aerofoil. That the central or partitioning wall may be formed integrally with the outer wall(s) of the aerofoil e.g. due to the manufacturing of the aerofoil by casting should not hinder this definition of these terms.
[0012] The feature that "the leading edge cavity is closed off in respect to an exterior" should be understood that an outer wall of the aerofoil positioned at the leading edge comprises no element, like a hole, an aperture, a gap, a slot or the like, that would allow flow communication between the leading edge cavity and an exterior of the aerofoil. Further, the feature, that the left and the right side cavities "are positioned at opposed sides of the leading edge cavity in respect to a fictitious (straight) connection of a midpoint of the at least one main impingement hole and the leading edge" should be understood in that, viewed in direction from the main impingement hole or the central wall to the leading edge, the left side cavity is positioned on the left hand side and the right side cavity at the right hand side of the leading edge cavity or the fictitious connection. Preferably, the side cavities are arranged in such a way in respect to the leading edge cavity so that the fictitious connection represents an axis of mirror symmetry.
[0013] Thus, the main cavity, the leading edge cavity and both side cavities build a channel system in flow communication with each other over the respective impingements holes. The main cavity is arranged upstream of the leading edge cavity as well as of the side cavities and the leading edge cavity, in turn, is positioned upstream of the side cavities, wherein the terms upstream and downstream refer to a flow direction of an flow of cooling medium and/or airflow through the turbine assembly. By this arrangement there is no direct flow communication between the side cavities.
[0014] The at least one main impingement hole and the side impingement holes may have any shape, size or orientation in the respective wall or positioning in the respective wall or towards each other feasible for a person skilled in the art and will be selected according to the cooling needs of the wall to be cooled or further characteristics of the turbine assembly (e.g. temperature or velocity of the cooling medium, pattern of the cooling passage, overall size, position in the engine). For example, the respective hole may be positioned centrically or off-centre in the respective wall e.g. for the central wall in reference to a direction from the suction side to the pressure side. Moreover, the respective hole may be arranged inclined or in normal direction in the wall.
[0015] Advantageously, the at least one main impingement hole is embodied in such a way so that a cooling medium is ejected in such a way so that it impinges an inner surface of the leading edge basically perpendicular in respect to the span-wise direction. Such an arrangement provided to be very effective for cooling the leading edge. In the scope of the impingement as "basically perpendicular" to an inner surface of the leading edge also lie a divergence of the impingement in respect to the inner surface of about 30°. Preferably, the impingement occurs perpendicular to the inner surface thus with an angle of 90°. In other words the cooling medium impinges basically with a direction that is in parallel to the direction from the leading edge to the trailing edge.
[0016] This may be realised by ejecting the cooling medium with a basically 90° angle from the main impingement hole. Therefore, the hole has (a) surrounding wall(s) with a basically, and preferably a strict, 90° angle in respect to a plane of the central wall extending in span-wise direction. In case the surrounding wall of the main impingement hole has also a basically, and preferably a strict, 90° angle in respect to a plane of the central wall extending perpendicular to the span-wise direction the cooling medium ejects from the main impingement hole perpendicular to the span-wise direction. The angle of the surrounding wall(s) may also be inclined in respect to the above mentioned planes of the central wall. The design of the holes (e.g. shape, size, orientation) depends on the needed cooling efficiency and the actual properties, like contour, material, size etc., of the aerofoil and will be selected by the person skilled in the art according to its knowledge in the field. All this features stated for the main impingement hole are also conferrable to the side impingement holes.
[0017] Preferably, the left side cavity is arranged directly at a pressure side of the hollow aerofoil and wherein the right side cavity is arranged directly at a suction side of the hollow aerofoil. Hence, transverse flow at the leading edge, which is the critical region because of the high heat transfer coefficients can be relegated to the pressure and suction sides where the external thermal loads are less important. Since the transverse flow can be avoided at the leading edge the most effective internal cooling system can take effect via the impingement cooling. In this context the phrase "arranged directly" should be understood as that the side cavities are arranged at an outer wall of the aerofoil or preferably that the outer wall of the aerofoil is also an outer wall of the respective side cavity. The outer wall of the left or right side cavity, respectively, may comprise a film cooling hole. But preferably, the side cavities are closed off in respect to the exterior of the hollow aerofoil.
[0018] In a further advantageous embodiment the side impingement holes have a cross sectional area that is smaller than a cross sectional area of the main impingement hole. Hence, jets discharged from the side impingement holes have a higher velocity flow that the jet from the main impingement hole. Consequently, the jets impinge effectively in the side cavities or their respective outer wall even if a cross radial flow is present in the side cavities or the lateral channels. Moreover, a reduced size of the side impingement holes compared with the main impingement hole avoids that the flow is directed in majority to the suction side hole instead to the pressure side hole e.g. in blades due to the rotating effect during operation. Hence, also a diameter of a side impingement hole is smaller than a diameter of the main impingement hole.
[0019] A ratio between the cross sectional area or size of the side impingement holes to the main impingement hole may have any value suitable for a person skilled in the art for example 0.1 to 0.9 (side impingement hole) to 1 (main impingement hole). Advantageously, the cross sectional area of a side impingement hole is about two times smaller than the cross sectional area of the main impingement hole (0.5/1). This size difference has been shown to be specifically advantageous. In this context the phrase "about two times" should be understood in that the size difference may have a ±10% deviation from the strict duplication of the size.
[0020] In an advantageous embodiment the at least one main impingement hole is positioned basically centrically in the least one shared central wall connecting the main cavity and the leading edge cavity. Thus, the jet discharged from the main impingement hole can be directed exactly at a leading edge point or in other words at the point of an inner surface of the outer wall that is the foremost point of the leading edge needing the most cooling efficiency. In this context "basically centrically" should be understood in that the positioning may have a ±10% deviation from the strict centric arrangement.
[0021] Beneficially, also each of the at least one side impingement holes is positioned basically centrically in the at least one partitioning wall providing effective cooling of the outer walls of the side cavities.
[0022] Preferably, the hollow aerofoil comprises at least one exit channel for the cooling medium extending along the hollow aerofoil from the leading edge to a trailing edge of the hollow aerofoil. Thus, the used cooling medium can be discharged effectively from the aerofoil. The exit channel may extend a whole distance from the leading edge to the trailing edge or only partially along the distance. Moreover, the exit channel may extend basically in parallel to the span-wise direction or angled thereto. Further, it discharges cooling medium via exit holed at the trailing edge of the aerofoil. The actual positioning of the exit channel depends on the construction or embodiment of the aerofoil. For example, is the aerofoil embodied as a blade the exit channel is advantageously positioned in a tip of the blade. In case the aerofoil is a vane the exit channel may be positioned in either the outer or inner platform, depending on where the vane is fed with cooling medium.
[0023] According to a further realisation of the invention the left side cavity and the right side cavity both have at least one exit aperture (left exit aperture and right exit aperture) and wherein the left side cavity and the right side cavity both discharge via the respective at least one exit aperture into the at least one exit channel (the left side cavity via the left exit aperture and the right side cavity via the right exit aperture). As a result, the cooling medium can be discharged directly. Hence, the side cavities connected the leading edge cavity with the at least one exit channel. There is no direct flow communication between the leading edge cavity and the exit channel. However, an inner wall of the exit channel may close off or cover and seal the leading edge cavity.
[0024] The at least one exit channel comprises at least one restricting wall being in respect to the exterior of the hollow aerofoil an outer wall of the hollow aerofoil. In other words the exit channel and the aerofoil share an outer wall. An even more direct discharge from the exit channel can be provided, when the at least one restricting wall of the at least one exit channel has at least one exit hole in its leading edge region. Moreover, this increases the flow area and consequently the mass flow of the cooling medium in the circuit and further provides cooling of a tip of the aerofoil. The exit hole has preferably a span-wise orientation, thus discharging the cooling medium e.g. at the tip of the blade.
[0025] In a preferred embodiment of the invention the at least one shared central wall connecting the main cavity and the leading edge cavity has a plurality of main impingement holes distributed along a span of the hollow aerofoil. Thus, the leading edge or the wall at the leading edge can be cooled effectively along the span of the aerofoil with different jets of cooling medium. In this context the phrase "along a span" should be understood as the extension of the aerofoil from its root to its tip or from one platform to the opposed platform. Moreover, the main impingement holes may be distributed along a part of the span or preferably along the whole span of the aerofoil. Furthermore, the distribution may have any pattern feasible for a person skilled in the art, like in pairs or groups or preferably evenly spaced apart along the span.
[0026] A number of side impingement holes can be any feasible for a person skilled in the art. This may depend on the dimension of the aerofoil or the respective cavities, respectively, or on the arrangement of the cavities towards each another or on the sizes of the impingement holes themselves.
[0027] Beneficially, each main impingement hole of the plurality of main impingement holes has a corresponding side impingement hole in the left partitioning wall and a corresponding side impingement hole in the right partitioning wall. In other words, each partitioning wall has the same amount of side impingement holes than the central wall has main impingement holes. Thus, the discharge of the cooling medium from the leading edge cavity to the side cavities can be realised homogeneously and directly.
[0028] Each side hole can be positioned at any span-wise height in respect to the respective main impingement hole and will be selected by a person skilled in the art according to the cooling needs. This may e.g. be necessary to balance the flow distribution in the leading edge channel, depending on the geometry of the cooling system.
[0029] The impinging action can be realised even more balanced when the side impingement holes are positioned in basically a same span-wise height as the main impingement hole. In this context "basically a same span-wise height" should be understood in that the positioning may have a ±10% deviation from the strict same span-wise height and "a same span-wise height" as being arranged on an axis extending perpendicular to the span-wise direction.
[0030] According to a further embodiment of the invention the leading edge cavity is divided in a plurality of sub-cavities along a span of the hollow aerofoil and wherein each sub-cavity comprises at least one main impingement hole and one side impingement hole to each side cavity. Due to this transverse flow can be avoided most efficiently. This is especially effective within blades to suppress the transverse flow due to the rotational effect of the blade during operation.
[0031] The ratio between the number of sub-cavities and the number of impingement holes (main/left side/right side) may be any suitable for a person skilled in the art, however preferred would be a 1:1:1:1 ratio. Moreover, the distribution and size of the sub-cavities may be any feasible for a person skilled in the art. Preferably, the sub-cavities are equally distributed and have basically the same size.
[0032] Furthermore, a turbulating structure may be arranged at any surface of the cavities. That may depend on the actual arrangement, shape, size, contour of the cavities or their respective walls or on the needed cooling and may be selected from the person skilled in the art according to its knowledge in the field. Moreover, such a turbulating structure may be any structure or element feasible for a person skilled in the art, like a recess, a projection, dimple, a turbulator or a fin.
[0033] According to a further aspect of the invention the left side cavity and/or the right side cavity comprise at least one inner surface of at least one wall of the side cavities and wherein at least parts of the inner surface comprise the at least one turbulating structure. Hence, the reduction of cooling efficiency of the impingement cooling due to the transverse flow in the side cavities can be effectively compensated. The phrase "at least parts of the inner surface" should be understood as sectionally or in sections. The wall may be any wall feasible for a person skilled in the art, preferably it is the wall at which the jet of cooling medium impinges and thus the outer wall of the side cavity and consequently of the aerofoil. An easy to realise turbulating structure can be provided when it is embodied as a dimple, a turbulator or a fin.
[0034] In a further realisation of the invention it is provided that the hollow aerofoil is a cast part out of one piece. Due to this impingement cooling can be advantageously used in blades without reducing the mechanical integrity for the blade. Thus, the aerofoil or the whole turbine assembly, respectively, is formed by casting during manufacturing of the turbine assembly.
[0035] In a further advantageous embodiment the aerofoil is a turbine blade or vane. Hence, an advantageous and effective cooling for these widely used components can be provided. This invention is combining impingement cooling without transverse flow at the leading edge location and optionally turbulated radial cavities or channels with impingement on the suction and pressure portion of the aerofoil leading edge. This configuration gives strength at the leading edge by using impingement cooling, which is the most effective internal cooling system if transverse flow is avoided. Moreover, the impingement cooling system is also realised in absence of film cooling holes which usually are used to drain the flow by reducing the transverse flow.
[0036] The above described and mentioned attributes, characteristics, features and advantages of this invention and the manner of attaining them will become more apparent and the invention itself will be better understood by reference to the following description of embodiments of the invention taken in conjunction with the accompanying drawings.
Brief Description of the Drawings
[0037] The present invention will be described with reference to drawings in which:
- FIG 1:
- shows a schematically and sectional view of a gas turbine engine comprising several inventive turbine assemblies,
- FIG 2:
- shows a cross section through a turbine assembly of FIG 1,
- FIG 3:
- shows a cross section through the turbine assembly along line III-III in FIG 2,
- FIG 4:
- shows schematically and simplified the flow path of a cooling medium traveling the turbine assembly from FIG 2 in span-wise direction,
- FIG 5
- shows a cross section through the schematically and simplified depicted turbine assembly along line V-V in FIG 4,
- FIG 6:
- shows a diagram depicting a heat transfer coefficients of different regions of the turbine assembly of FIG 2 and
- FIG 7:
- shows an alternative turbine assembly embodied as a vane.
Detailed Description of the Illustrated Embodiments
[0038] The present invention is described with reference to an exemplary turbine engine 82 having a single shaft 94 or spool connecting a single, multi-stage compressor section 86 and a single, one or more stage turbine section 90. However, it should be appreciated that the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
[0039] The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine 82 unless otherwise stated. If used, the terms axial, radial and circumferential are made with reference to a rotational axis 92 of the engine 82.
[0040] FIG 1 shows an example of a gas turbine engine 82 in a sectional view. The gas turbine engine 82 comprises, in flow series, an inlet 84, a compressor section 86, a combustion section 88 and a turbine section 90, which are generally arranged in flow series and generally in the direction of a longitudinal or rotational axis 92. The gas turbine engine 82 further comprises a shaft 94 which is rotatable about the rotational axis 92 and which extends longitudinally through the gas turbine engine 82. The shaft 94 drivingly connects the turbine section 90 to the compressor section 86.
[0041] In operation of the gas turbine engine 82, air 96, which is taken in through the air inlet 84 is compressed by the compressor section 86 and delivered to the combustion section or burner section 88. The burner section 88 comprises a burner plenum 98, one or more combustion chambers 100 defined by a double wall can 102 and at least one burner 104 fixed to each combustion chamber 100. The combustion chambers 100 and the burners 104 are located inside the burner plenum 98. The compressed air passing through the compressor section 88 enters a diffuser 106 and is discharged from the diffuser 106 into the burner plenum 98 from where a portion of the air enters the burner 104 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 108 or working gas from the combustion is channelled via a transition duct 110 to the turbine section 90.
[0042] This exemplary gas turbine engine 82 has a cannular combustor section arrangement 112, which is constituted by an annular array of combustor cans 102 each having the burner 104 and the combustion chamber 100, the transition duct 110 has a generally circular inlet that interfaces with the combustion chamber 100 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 90.
[0043] The turbine section 90 comprises a number of blade carrying discs 114 or turbine wheels 116 attached to the shaft 94. In the present example, the turbine section 90 comprises two discs 114 each carry an annular array of turbine assemblies 10, which each comprises an aerofoil 12 embodied as a turbine blade 78. However, the number of blade carrying discs 114 could be different, i.e. only one disc 114 or more than two discs 114. In addition, turbine cascades 118 are disposed between the turbine blades 78. Each turbine cascade 118 carries an annular array of turbine assemblies 10a, which each comprises an aerofoil 12 in the form of guiding vanes 80, which are fixed to a stator 120 of the gas turbine engine 82. Between the exit of the combustion chamber 100 and the leading turbine blades 78 inlet guiding vanes or nozzle guide vanes 122 are provided and turn the flow of working gas 108 onto the turbine blades 78.
[0044] The combustion gas 108 from the combustion chamber 100 enters the turbine section 90 and drives the turbine blades 78 which in turn rotate the shaft 94. The guiding vanes 122 serve to optimise the angle of the combustion or working gas 108 on to the turbine blades 78. The turbine section 90 drives the compressor section 86. The compressor section 86 comprises an axial series of guide vane stages 124 and rotor blade stages 126. The rotor blade stages 126 comprise a rotor disc 114 supporting turbine assemblies 10 with an annular array of aerofoils 12 or turbine blades 78.
[0045] The compressor section 124 also comprises a stationary casing 128 that surrounds the rotor stages 126 in circumferential direction 130 and supports the vane stages 124. The guide vane stages 124 include an annular array of radially extending turbine assemblies 10a with aerofoils 12 embodied as vanes 80 that are mounted to the casing 128. The vanes 80 are provided to present gas flow at an optimal angle for the blades 78 at a given engine operational point. Some of the guide vane stages 124 have variable vanes 80, where the angle of the vanes 80, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.
[0046] The casing 128 defines a radially outer surface 132 of a passage 134 of the compressor section 86. A radially inner surface 136 of the passage 134 is at least partly defined by a rotor drum 138 of the rotor which is partly defined by the annular array of blades 78.
[0047] FIG 2 shows a cross section through a turbine assembly 10 of the gas turbine engine 82. The turbine assembly 10 comprises a basically hollow aerofoil 12, embodied as a turbine blade 78, with two cooling regions, specifically, an impingement cooling region 140 and a fin-pin/pedestal cooling region 142. The former is located at a leading edge 18 and the latter at a trailing edge 54 of the aerofoil 12. At opposed ends 144, 144' the aerofoil 12 comprises a root portion 146 and a tip 148.
[0048] The aerofoil 12, which is cast during the manufacturing of the turbine assembly 10 as a cast part out of one piece, comprises a casing 150 that forms several internal chambers or passages. Specifically, two feeding channels 152 positioned in a leading edge region 66 of the aerofoil 12, extending through the root portion 146 and feeding cooling medium 44, like air, to a main cavity 14 and to a maze-like or race track passage 154 meandering to the trailing edge 54. The main cavity 14 is positioned in the leading edge region 66 and is spanning the aerofoil 12 in span-wise direction 20 basically from the root 146 to the tip 148.
[0049] Furthermore, the aerofoil 12 comprises a leading edge cavity 16 positioned directly at the leading edge 18 that also extends in span-wise direction 20. Moreover, the leading edge cavity 16 is closed off in respect to an exterior 26 of the aerofoil 12; in other words, an outer wall 156 of the leading edge cavity 16 that is also an outer wall 62 of the aerofoil 12 at the leading edge 18 has no openings, like film cooling holes. For connecting the main cavity 14 and the leading edge cavity 16 with one another the aerofoil 12 comprises a shared central wall 22 extending in direction 162 pointing from a pressure side 48 to a suction side 50 or vice versa.
[0050] To provide sufficient and effective cooling of the leading edge 18 or its outer wall 156 the central wall 22 comprises a plurality of main impingement holes 24 distributed along a span 68 of the aerofoil 12. The main impingement holes 24 are evenly distributed in span-wise direction 20 along the central wall 22 and each one is positioned basically centrically in the shared central wall 22 in direction 162 perpendicular to the span-wise direction 20. This can be seen in in FIG 3 that shows a cross section of the turbine assembly 10 along line III-III in FIG 2.
[0051] To suppress transverse flow especially enhanced by the rotation of the blade 78 during operation the leading edge cavity 16 is divided in a plurality of sub-cavities 70 along the span 68 of the aerofoil 12 (see FIG 2). Therefore, partition walls 158 are arranged equally spaced apart along the span 68. The main cavity 14 is connected with each sub-cavity 70 via one main impingement hole 24.
[0052] Each main impingement hole 24 is embodied in such a way so that the cooling medium 44 is ejected in such a way so that it impinges an inner surface 46 of the leading edge 18 basically perpendicular in respect to the span-wise direction 20. Therefore, a surrounding wall 160 of the main impingement hole 24 has an angle of 90° in respect to extensions of the central wall 22 in span wise direction 20 and in direction 162 perpendicular to the span wise direction 20.
[0053] Due to the central positioning in the central wall 22 and the 90° angle of the surrounding wall 160 of the main impingement hole 24 jets of cooling medium 44 that are discharged from the main impingement holes 24 are directed directly at the inner surface 46 of the outer wall 156 of the leading edge 18 or specifically to a leading edge point 164 which is the point of the inner surface 46 of the outer wall 156 that is the foremost point of the leading edge 18 and thus is needing the most cooling efficiency.
[0054] Further, the aerofoil 12 comprises a left side cavity 28 and a right side cavity 30 that each extent in span-wise direction 20. The left side cavity 28 is connected to the leading edge cavity 16 by a shared left partitioning wall 32 and the right side cavity 30 is connected to the leading edge cavity 16 by a shared right partitioning wall 34. Moreover, the left partitioning wall 32 and the right partitioning wall 34 are positioned at opposed sides (left and right) of the leading edge cavity 16 in respect to a fictitious straight connection 36 of a midpoint 38 of each main impingement hole 24 and the leading edge 18 or its respective leading edge point 164. Specifically, the left side cavity 28 is arranged directly at the pressure side 48 of the aerofoil 12 and the right side cavity 30 is arranged directly at the suction side 50 of the aerofoil 12.
[0055] For discharge of the cooling medium 44 from the leading edge cavity 16 each of the partitioning walls 32, 34 comprise a plurality of left and right side impingement holes 40, 42. Each side impingement hole 40, 42 is positioned basically centrically in the respective left or right partitioning wall 32, 34. Each side impingement hole 40, 42 is embodied in such a way so that the cooling medium 44 is ejected in such a way so that it impinges an inner surface 72 of an outer wall 74 of the side cavities 28, 30, wherein the outer wall 74 is also the outer wall 62 of the aerofoil 12 positioned at the pressure side 48 or the suction side 50, respectively. Preferably, a surrounding wall 160 of the side impingement holes 40, 42 have also a 90° angle in respect to extensions of the respective partitioning wall 32, 34. However, different angles are also possible.
[0056] Each main impingement hole 24 has a corresponding side impingement hole 40 in the left partitioning wall 32 and a corresponding side impingement hole 42 in the right partitioning wall 34. Further, the side impingement holes 40, 42 are positioned in basically a same span-wise height h as the main impingement hole 24. Thus, each sub-cavity 70 comprises at least one main impingement hole 24 and a left and right side impingement hole 40, 42 connecting the sub-cavity 70 to each side cavity 28, 30.
[0057] In this exemplary embodiment the main and side impingement holes 24, 40, 42 are positioned centrically in respect to two (an upper and a lower) partition walls 158. It may be also feasible and even advantageous to position the main and side impingement holes at different heights or at different positions in respect to the respective partition wall. E.g. to position the main impingement hole close to the lower partition wall and the side impingement holes close to the upper partition wall or to position the main impingement hole centrically and the side impingement holes close to the upper partition wall (not shown).
[0058] To allow the cooling medium 44 to exit the aerofoil 12 it comprises an exit channel 52 extending along the aerofoil 12 from the leading edge 18 to the trailing edge 54 and in this exemplary embodiment in the tip 148 of the aerofoil 12. To communicate with the exit channel 52 the left side cavity 28 and the right side cavity 30 both have at least one exit aperture 56, 58 that both discharge via the respective exit aperture 56, 58 into the exit channel 52.
[0059] Moreover, the exit channel 52 comprises a restricting wall 60 being in respect to the exterior 26 of the aerofoil 12 an outer wall 62 of the aerofoil 12 and thus being arranged at the tip 148 of the aerofoil 12. For providing an additional discharge of the cooling medium 44 the restricting wall 60 has exit holes 64 in its leading edge region 66.
[0060] Parts of the inner surface 72 of the wall 74 (outer wall 62) of the left and right side cavities 28, 30 comprise a turbulating structure 76 that is for example embodied as a dimple, a turbulator or a fin.
[0061] Moreover, the side impingement holes 40, 42 have a cross sectional area a and a dimeter d that is smaller than a cross sectional area A and a diameter D of the main impingement hole 24 and specifically the cross sectional area a of a side impingement hole 40, 42 is about two times smaller than the cross sectional area A of the main impingement hole 24 (see also FIG 5). Consequently, the velocity of the cooling medium 44 being ejected from the side impingement holes 40, 42 is higher than the velocity of the cooling medium 44 being ejected from the main impingement hole 24. The higher velocity can compensate the transverse flow occurring in the side cavities 28, 30.
[0062] The differences in area a, A and diameter d, D of the main and side impingement holes 24, 40, 42 is depicted in FIG 4 that shows schematically and simplified the flow path of the cooling medium 44 traveling the turbine assembly 10 in span-wise direction 20 and in FIG 5 that shows a cross section through the schematically and simplified depicted turbine assembly 10 along line V-V in FIG 4.
[0063] The flow path of the cooling medium 44 from the main cavity 14 via the main impingement holes 24, the leading edge cavity 16, the side impingement holes 40, 42, the side cavities 28, 30 and the exit channel 52 towards the trailing edge 54 can be seen in FIG 4 and 5 (for better presentability the flow is not shown for all impingement holes 24, 40, 42).
[0064] As can be seen in FIG 4 and 5 transverse flow occurs in the side cavities 28, 30. That may reduce the cooling effect of the impingement cooling. However, this can be neglected since the external thermal loads at these regions of the pressure and suction side 48, 50 are less critical than at the leading edge 18. This can be seen in FIG 6 that shows a diagram depicting a heat transfer coefficients (y-axis HTC) of different regions of an aerofoil profile cross section of the turbine assembly 10 (pressure side 48, leading edge 18, suction side 50) (x-axis as S/Smax; a distance (S) from the trailing edge divided by the total distance (Smax) in the aerofoil profile section).
[0065] In FIG 7 an alternative embodiment of the turbine assembly is shown. Components, features and functions that remain identical are in principle substantially denoted by the same reference characters. To distinguish between the embodiments, however, the letter "a" has been added to the different reference characters of the embodiment in FIG 1 to 6. The following description is confined substantially to the differences from the embodiment in FIG 1 to 6, wherein with regard to components, features and functions that remain identical reference may be made to the description of the embodiment in FIG 1 to 6.
[0066] FIG 7 shows an alternative turbine assembly 10a that is embodied as a vane 80. The embodiment from FIG 7 differs in regard to the embodiment according to FIG 1 to 6 in that a leading edge cavity 16a at a leading edge 18 is a single chamber extending in span-wise direction 20 undivided along a span 68 of an aerofoil 12 of the turbine assembly 10a. Since a vane 80 is a stationary element in a turbine engine 82 the main driver for a flow of a cooling medium 44 is the pressure difference between an inlet and an exit of the medium 44. Hence, since there are not centrifugal forces a transverse flow can be neglected. In fact the flow coming from the main impingement holes 24 is leaving the leading edge cavity 16a through the side impingement holes 40, 42 closest to each main impingement hole 24. This is due to the principle of the minimum distance for the cooling medium 44 to move from the region at high pressure (main cavity 14) to the region at lower pressure (side cavities 28, 30).
[0067] It should be noted that the term "comprising" does not exclude other elements or steps and "a" or "an" does not exclude a plurality. Also elements described in association with different embodiments may be combined. It should also be noted that reference signs in the claims should not be construed as limiting the scope of the claims.
1. A turbine assembly (10, 10a) comprising a basically hollow aerofoil (12) having at
least a main cavity (14) and a leading edge cavity (16, 16a) positioned at a leading
edge (18) of the hollow aerofoil (12), wherein the main cavity (14) and the leading
edge cavity (16, 16a) extent at least partially in span-wise direction (20) of the
hollow aerofoil (12) and wherein the hollow aerofoil (12) comprises at least one shared
central wall (22) connecting the main cavity (14) and the leading edge cavity (16,
16a), wherein the at least one central wall (22) comprises at least one main impingement
hole (24), and wherein the leading edge cavity (16, 16a) is closed off in respect
to an exterior (26) of the hollow aerofoil (12), characterized in that the hollow aerofoil (12) comprises at least a left side cavity (28) and a right side
cavity (30), wherein each side cavity (28, 30) extent at least partially in span-wise
direction (20) of the hollow aerofoil (12) and wherein the left side cavity (28) is
connected to the leading edge cavity (16, 16a) by at least one shared left partitioning
wall (32) and the right side cavity (30) is connected to the leading edge cavity (16,
16a) by at least one shared right partitioning wall (34), wherein the at least one
left partitioning wall (32) and the at least one right partitioning wall (34) are
positioned at opposed sides of the leading edge cavity (16, 16a) in respect to a hypothetical
connection (36) of a midpoint (38) of the at least one main impingement hole (24)
and the leading edge (18) and wherein each of the at least one partitioning walls
(32, 34) comprise at least one side impingement hole (40, 42).
2. A turbine assembly according to claim 1, wherein the at least one main impingement
hole (24) is embodied in such a way so that a cooling medium (44) is ejected in such
a way so that it impinges an inner surface (46) of the leading edge (18) basically
perpendicular in respect to the span-wise direction (20).
3. A turbine assembly according to claim 1 or 2, wherein the left side cavity (28) is
arranged directly at a pressure side (48) of the hollow aerofoil (12) and wherein
the right side cavity (30) is arranged directly at a suction side (50) of the hollow
aerofoil (12).
4. A turbine assembly according to any preceding claim, wherein the side impingement
holes (40, 42) have a cross sectional area (a) that is smaller than a cross sectional
area (A) of the main impingement hole (24), specifically, wherein the cross sectional
area (a) of a side impingement hole (40, 42) is about two times smaller than the cross
sectional area (A) of the main impingement hole (24).
5. A turbine assembly according to any preceding claim, wherein the at least one main
impingement hole (24) is positioned basically centrically in the least one shared
central wall (22) connecting the main cavity (14) and the leading edge cavity (16,
16a) and/or wherein each of the at least one side impingement holes (40, 42) is positioned
basically centrically in the at least one partitioning wall (32, 34).
6. A turbine assembly according to any preceding claim, wherein the hollow aerofoil (12)
comprises at least one exit channel (52) extending along the hollow aerofoil (12)
from the leading edge (18) to a trailing edge (54) of the hollow aerofoil (12) and
wherein the left side cavity (28) and the right side cavity (30) both have at least
one exit aperture (56, 58) and wherein the left side cavity (28) and the right side
cavity (30) both discharge via the respective at least one exit aperture (56, 58)
into the at least one exit channel (52).
7. A turbine assembly according to claim 6, wherein the at least one exit channel (52)
comprises at least one restricting wall (60) being in respect to the exterior (26)
of the hollow aerofoil (12) an outer wall (62) of the hollow aerofoil (12) and wherein
the at least one restricting wall (60) has at least one exit hole (64) in its leading
edge region (66).
8. A turbine assembly according to according to any preceding claim, wherein the at least
one shared central wall (22) connecting the main cavity (14) and the leading edge
cavity (16, 16a) has a plurality of main impingement holes (24) distributed along
a span (68) of the hollow aerofoil (12).
9. A turbine assembly according to any preceding claim, wherein the at least one shared
central wall (22) connecting the main cavity (14) and the leading edge cavity (16,
16a) has a plurality of main impingement holes (24) and wherein each main impingement
hole (24) of the plurality of main impingement holes (24) has a corresponding side
impingement hole (40) in the left partitioning wall (32) and a corresponding side
impingement hole (42) in the right partitioning wall (34), wherein the side impingement
holes (40, 42) are positioned in basically a same span-wise height (h) as the main
impingement hole (24).
10. A turbine assembly according to any preceding claim, wherein the leading edge cavity
(16) is divided in a plurality of sub-cavities (70) along a span (68) of the hollow
aerofoil (12) and wherein each sub-cavity (70) comprises at least one main impingement
hole (24) and one side impingement hole (40, 42) to each side cavity (28, 30).
11. A turbine assembly according to according to any preceding claim, wherein the left
side cavity (28) and/or the right side cavity (30) comprise at least one inner surface
(72) of at least one wall (74) of the side cavities (28, 30) and wherein at least
parts of the inner surface (72) comprise at least one turbulating structure (76),
especially, the at least one turbulating structure is embodied as a dimple, a turbulator
or a fin.
12. A turbine assembly according to according to any preceding claim, wherein the hollow
aerofoil (12) is a cast part out of one piece.
13. A turbine assembly according to any preceding claim, wherein the hollow aerofoil (12)
is a turbine blade (78) or vane (80).