• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
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  • 优先权
    • 专利摘要:
    The invention relates to a blade (16) for a turbine of a turbomachine, which comprises a blade (25), wherein the blade (25) has a leading edge (28), a trailing edge (29), an outer tip and an inner end. The blade (25) further includes a cooling structure having a plurality of cooling channels (33) for receiving and directing a coolant. The plurality of cooling channels (33) may include a linear cooling channel and a curved cooling channel (34). The blade (16) also has a contoured shape defined by the blade (25) between the inboard end and the outboard tip, the contoured shape being configured to include a target area (41) that is defined by a linear datum line, which extends radially from a location at the inboard end of the blade (25), can not be achieved. The curved cooling passages (34) are configured to extend between an upstream end and a downstream end so as to intersect the target area (41) therebetween.
    公开号:CH710576A2
    申请号:CH01854/15
    申请日:2015-12-17
    公开日:2016-06-30
    发明作者:Anthony Weber Joseph；James Healy Michael
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    BACKGROUND OF THE INVENTION
    The present invention relates to internal cooling channels of turbine blades in gas turbines. More specifically, but not by way of limitation, the present invention relates to non-linear internal cooling channels formed within long blades within the aft rows of turbine blades.
    Note that gas turbines generally include a compressor, a combustor, and a turbine. The compressor and turbine sections generally have rows of blades staggered axially in steps. Each stage has a series of circumferentially spaced vanes that are fixed and a series of blades that rotate about a central turbine axis or shaft. In operation, the compressor blades rotate about the shaft and cooperate with the vanes to compress an airflow. The supplied compressed air is then used in the combustion chamber to burn supplied fuel. The resulting stream of hot, expanding gases from combustion, i. the working fluid expands on its way through the turbine section of the engine. The flow of working fluid through the turbine causes the blades to rotate. The blades are connected to a central shaft such that the rotation of the blades rotates the shaft.
    In this way, the energy contained in the fuel is converted into the mechanical energy of the rotary shaft, which can be used, for example, to rotate the blades of the compressor, so that the supply of compressed air required for combustion is produced , as well as the coils of a generator, so that electrical power is generated. During operation, the turbine blades, which as described generally include both the rotating blades and the fixed circumferentially-spaced vanes, are heavily stressed by extreme mechanical and thermal stresses due to the extreme temperatures of the hot gas path, the velocity of the working fluid, and the speed of the machine ,
    Because of the ever-increasing energy demand continues and urgently strives to develop more efficient gas turbines. Although several strategies for increasing the efficiency of turbomachines are known, this remains a difficult-to-achieve goal, as these alternatives, which include, for example, increasing the size of the engine, increasing the temperatures in the hot gas path, and increasing the number of revolutions of the blades, generally represent additional load on the parts that are already heavily loaded, such as the blades and vanes of the turbine. As a result, there is an urgent need for improved apparatus, methods and / or systems that reduce operational loads on the turbine blades or that allow the turbine blades to better withstand these loads. One of ordinary skill in the art will recognize that one strategy for reducing the heat load on the blades is to cool them during operation. Effective cooling, for example, may allow the blades to withstand higher firing temperatures, withstand higher mechanical stresses at high operating temperatures, and / or extend the life of blade components, all allowing for more efficient and efficient operation of the turbomachine. One way to cool the blades during operation is to use internal coolant passageways or circuits. Generally, this involves passing a supply of relatively cool compressed air, which may be supplied by the compressor of the turbomachine, through internal cooling channels within the blades. As the compressed air flows through the blade, it cools the blade by convection, allowing the component to withstand burning temperatures that otherwise would not endure.
    It will be appreciated that great care is needed in designing the structure of these cooling channels during their manufacture for several reasons. First, the use of cooling air has an impact. That is, air that is diverted from the compressor to the turbine section of the engine for cooling bypasses the combustion chamber and thus reduces the efficiency of the engine. Thus, coolant passageways must be designed so that air is used in a very efficient manner, i. the necessary surface supply and cooling effect is provided, so that only the smallest possible amount of air is required for this purpose. Second, newer, more aggressively shaped aerodynamic blade structures are thinner and more curved or tortuous, which often makes it impossible to use linear cooling channels that run the entire length of the turbine blade, while the small thickness of the blades requires high performance of the coolant conduit paths that still require have to have a compact design. Third, to reduce mechanical stresses, coolant passageways may be formed to remove excess weight from the blade; However, the blades must still be strong to withstand the extreme mechanical loads. Cooling channels must therefore be designed so that the turbine blade has a light but strong construction, while avoiding stress concentrations that would adversely affect the load capacity of the blades. Thus, there is a commercial need for turbine blade cooling structures that are well suited for more aggressive shaped, thinner aerodynamic blade structures, favoring a lighter inner design of the blade, retaining structural strength of the component, and thereby providing high cooling efficiency.
    BRIEF DESCRIPTION OF THE INVENTION
    The present invention thus describes a blade for a turbine of a gas turbine having a blade, wherein the blade has a leading edge, a trailing edge, an outer tip and an inner end, wherein the sheet is attached to a foot, the purpose is designed to connect the turbine blade with a rotor disk. The blade may further include a cooling structure that includes a plurality of elongated cooling channels for receiving and directing coolant through the blade. The plurality of cooling channels may include at least one linear cooling channel and a curved cooling channel. The blade may also have a contoured shape defined by the blade between the inboard end and the outboard tip, wherein the contoured shape is configured to include a target area extending from a linear datum line that extends radially from a location at the inboard end of the blade Leaf runs out, can not be reached. The curved cooling passages may be configured to extend between a downstream end and a downstream end so as to intersect the target area therebetween.
    In any embodiment of the invention, it may be advantageous if the coolant inlet extends through the foot of the blade, so that it is connected to a coolant source; wherein the blade comprises a blade and the curved cooling channel has a feature formed after casting; and wherein the linear reference line includes one extending between the outer tip and the inner end of the blade.
    [0008] In any embodiment of the invention, it may be advantageous for the curved coolant passageway at the upstream end to communicate with a downstream end of a linear refrigerant passageway, the connection comprising a position within the blade; wherein the linear refrigerant passage extends between the connection with the curved coolant passage and a connection with the coolant inlet.
    In any embodiment of the invention, it may be advantageous for the contoured shape of the sheet to comprise at least one of: a concave-shaped pressure-side surface and a convex-shaped suction-side surface joined along the upstream and downstream edges; and a radial diffraction component with which an arc is formed along a longitudinal axis of the sheet.
    In any embodiment of the invention, it may be advantageous that the contoured shape of the sheet has a twist about a longitudinal axis of the sheet, wherein the distortion is designed so that a staking angle for the sheet generally between the inner end and the outside tip varies.
    In any embodiment of the invention, it may be advantageous for the contoured shape of the blade to have a taper along a longitudinal axis of the blade, the taper comprising at least one of the following: an axial taper through which a distance between the leading edge and the trailing edge between the inner end and the outer tip of the sheet is gradually smaller; a taper in the circumferential direction by which a thickness between the pressure-side surface and the suction-side surface between the inner end and the outer tip of the sheet is gradually reduced.
    In any embodiment of the invention, it may be advantageous for the structure of the linear inaccessibility of the target area to comprise a combination of at least two of: the twist about the longitudinal axis of the blade; the taper along the longitudinal axis of the blade; and the radial diffraction component by which the arc is defined along the longitudinal axis of the sheet; and wherein a curvature of the curved coolant passageway is shaped to correspond to a curvature of the contoured shape of the sheet.
    [0013] In any embodiment of the invention, it may be advantageous for the downstream end of the curved coolant conduit path to be connected to an exit port through an outer surface of the blade, the exit port being located at least one of: the outer tip; pressure side surface; suction surface; leading edge; Trailing edge.
    In any embodiment of the invention, it may be advantageous for the curved coolant conduit path to have a connection with the coolant inlet at the upstream end, the connection comprising a position near the inboard end of the blade.
    [0015] In any embodiment of the invention, it may be advantageous for the curved coolant passageway at the upstream end to communicate with a downstream end of a linear refrigerant passageway, the connection comprising a position within the blade; wherein the linear refrigerant passage extends between the connection with the curved coolant passage and a connection with the coolant inlet.
    In any embodiment of the invention, it may be advantageous for the blade to have a leading half / leading half and a tracked half / tail half defined on either side of an axial centerline connecting centers of the sheet camber lines; the blade comprising radially staggered portions defined within and out of a radial centerline of the blade, an inboard portion extending between the root and the radial centerline and an outboard portion extending between the radial centerline and the outboard tip; and wherein the cooling structure comprises a plurality of curved cooling channels.
    In any embodiment of the invention, it may be advantageous for the upstream end of each curved cooling channel to include a position near the axial centerline and the inboard end of the blade; and wherein the curved cooling channels include a curve that is curved toward the outlet openings formed on at least one of the leading edge of the sheet; Trailing edge of the sheet.
    In any embodiment of the invention, it may be advantageous for each of the curved cooling channels to extend parallel with respect to the other curved cooling channels and along a curvature curved to the trailing edge in the sheet; wherein the outlet openings comprise a radial distance along the trailing edge of the sheet.
    In any embodiment of the invention, it may be advantageous that the upstream ends of the curved cooling channels are positioned within the leading half of the sheet and the outlet openings are formed through the outer surface of the trailing half of the sheet.
    In any embodiment of the invention, it may be advantageous for the blade to be configured to operate in a multi-stage turbine in a rear row of blades; wherein the target area is positioned within the outboard section of the sheet.
    In any embodiment of the invention, it may be advantageous for the target area to be located within the tracked half of the sheet, and for each of the curved cooling channels to include a curvature curved toward the trailing edge of the sheet.
    In any embodiment of the invention, it may be advantageous that the cooling structure comprises a plurality of the linear cooling channels and a plurality of curved cooling channels, each having a downstream end in the vicinity of the inner end of the sheet; and each of the linear cooling channels and the curved cooling channels extends over at least a majority of a radial height of the sheet.
    In any embodiment of the invention, it may be advantageous that each of the curved cooling channels extends in parallel from an inboard position, where they are each connected at a central junction with one of the linear cooling channels, to an outboard position where they are each connected to one of the outlet openings.
    In any embodiment of the invention it may be advantageous that one of the curved cooling channels has a curvature corresponding to a surface contour of the pressure-side surface of the sheet; and wherein each of the curved cooling channels includes a curvature corresponding to a surface contour of the suction-side surface of the sheet.
    In any embodiment of the invention, it may be advantageous for the curved cooling channels to extend radially with respect to a curvature of a portion of the outer surface of the sheet; and wherein the curved cooling channel maintains a substantially constant distance from the area of the outer surface of the sheet.
    In any embodiment of the invention, it may be advantageous for the curved cooling channels to comprise a feature formed after casting which is formed after the casting of the blade by a removal method; and wherein the ablation process comprises a steerable electrochemical ablation process or an electrical discharge based ablation process.
    These and other features of the present application will become apparent upon consideration of the following detailed description of the preferred embodiments, taken in conjunction with the drawings and the appended claims.
    BRIEF DESCRIPTION OF THE DRAWINGS
    These and other features of the invention will be better understood by reading the following more detailed description of embodiments of the invention in conjunction with the accompanying drawings, in which:<Tb> FIG. 1 <SEP> is a schematic representation of an example of a turbomachine in which blades according to embodiments of the present invention can be used;<Tb> FIG. FIG. 2 is a cross-sectional view of the compressor section of the combustion combustion engine of FIG. 1; FIG.<Tb> FIG. FIG. 3 is a cross-sectional view of the turbine section of the combustion combustion engine of FIG. 1; FIG.<Tb> FIG. 4 is a side view of an example of a turbine blade in which embodiments of the present invention may be used;<Tb> FIG. Fig. 5 is a cross-sectional view taken along line 5-5 of Fig. 4;<Tb> FIG. Fig. 6 <SEP> is a cross-sectional view taken along line 6-6 of Fig. 4;<Tb> FIG. Fig. 7 is a cross-sectional view taken along the line 7-7 of Fig. 4;<Tb> FIG. Fig. 8 is a cross-sectional view taken along line 8-8 of Fig. 4;<Tb> FIG. 9 is a perspective view of an example of a turbine blade having a twisted, curved, and tapered structure in which embodiments of the present invention may be used;<Tb> FIG. Fig. 10 <SEP> is a partial cross-sectional view taken along line 10-10 of Fig. 9;<Tb> FIG. Fig. 11 is a side view of a turbine blade illustrating conventional linear cooling channels;<Tb> FIG. Fig. 12 is a perspective view of the turbine blade illustrated in Fig. 11;<Tb> FIG. FIG. 13 is a side view of a turbine blade illustrating curved cooling channels according to one embodiment of the present invention; FIG.<Tb> FIG. FIG. 14 is a perspective view of the turbine blade illustrated in FIG. 13; FIG.<Tb> FIG. Fig. 15 illustrates a perspective view of a turbine bucket with curved cooling channels according to an alternative embodiment of the present invention;<Tb> FIG. 16 is a perspective view of a turbine blade illustrating curved cooling channels according to an alternative embodiment of the present invention; and<Tb> FIG. 17 illustrates a perspective view of a turbine blade with curved cooling channels according to an alternative embodiment of the present invention.
    DETAILED DESCRIPTION OF THE INVENTION
    Aspects and advantages of the invention will be set forth in the description which follows, or may be learned from the description, or may be learned by practice of the invention. Reference will now be made in detail to present embodiments of the invention, for which one or more examples are illustrated in the accompanying drawings. The detailed description numerical designations to refer to features in the drawings. The same or similar terms in the drawings and the description may be used to refer to the same or similar parts of the embodiments of the invention. Note that each example is given to illustrate the invention, but not for the purpose of limiting the invention. In fact, those skilled in the art will recognize that modifications and changes may be made to the present invention without departing from its scope or spirit. For example, features illustrated or described as part of one embodiment may be used in another embodiment to yield still another embodiment. Thus, the present invention is intended to cover such modifications and changes that come within the scope of the appended claims and their equivalents. Note that the ranges and limits referred to herein include all portions that are within the given limits, including limits per se, unless otherwise specified. In addition, certain terms have been chosen to describe the present invention and the subsystems and components of which it is constructed. As far as possible, these terms have been chosen on the basis of terminology, which is common in the art. Nevertheless, it should be noted that such terms are often interpreted differently. For example, what may be referred to herein as a single component may be considered elsewhere as consisting of multiple components, or something referred to herein as comprising multiple components may be considered as a single component elsewhere. In order to understand the scope of the present invention, one should not only pay attention to the particular terminology used, but also to the accompanying description and context, as well as the structure, design, function and / or use of the named and described component, including the manner in which the term refers to the several figures, as well as, of course, to the exact use of the terminology in the appended claims. Although the following examples are presented with respect to a particular type of turbine, the technique of the present invention may be applicable to other types of turbines, as one of ordinary skill in the relevant art will recognize.
    In view of the nature of turbomachine operation, several illustrative terms may be used in the specification to explain the operation of the machine and / or the multiple subsystems or components contained therein, and it may be advantageous to use these terms at the beginning of this section define. Thus, unless otherwise indicated, these terms and their definitions are as follows. Without further specification, the terms "front" and "rear" refer to directions with respect to the orientation of the gas turbine. That is, "front" refers to the front or compressor end of the engine, and "rear" refers to the rear or turbine end of the engine. Note that each of these terms can be used to indicate movement or relative position within the machine. The terms "downstream" and "upstream" are used to indicate a position within a particular conduit relative to the general direction of the flow therethrough. (Note that these terms indicate direction with respect to expected current during normal operation, which will be understood by one of ordinary skill in the art). The term "downstream" refers to the direction in which the fluid flows through the designated conduit, while "upstream" refers to the opposite direction. For example, the primary flow of a working fluid through a turbomachine initially consisting of air that moves through the compressor and then becomes fuel gases in the combustion chamber and beyond it may be described as flowing at an upstream direction Point begins near an upper or front end of the compressor and ends at a downstream point in the vicinity of a lower or rear end of the turbine. As regards the description of a flow within a combustion chamber of a conventional type, it should be noted that, as will be discussed in more detail below, air discharged from the compressor typically enters the combustion chamber through baffles which (in relation to the longitudinal axis of the combustion chamber and combustion chamber) the above-mentioned compressor / turbine positions defining distinctions between front / rear) are concentrated to the rear end of the combustion chamber. Upon entering the combustion chamber, the compressed air is directed from an annular passage formed around an inner chamber to the forward end of the combustion chamber where the air flow enters the inner chamber and, after having reversed its flow direction, to the rear end of the combustion chamber flows. In yet another context, coolant flows through coolant passageways may be treated in the same manner.
    Moreover, in view of the fact that the compressor and the turbine are designed with a common center axis, as well as the cylindrical structure common to many types of combustors, terms describing a position with respect to an axis may be used herein. As far as this is concerned, it should be noted that the term "radial" refers to a movement or position perpendicular to an axis. In this connection, it may be necessary to describe a relative distance from the central axis. For example, if a first component is closer to the central axis than a second component, then in this case the first component is described as "radially inward" or "inward" of the second component. Conversely, if the first component is farther away from the central axis than the second component, the first component is described herein as "radially outboard" or "outward" of the second component. It should also be noted that the term "axial" refers to a movement or position parallel to an axis. Finally, the term "in the circumferential direction" refers to a movement or position about an axis. While these terms may, as stated, be used with respect to the common centerline extending through the compressor and turbine sections of the machine, these terms may also be used in relation to other components or subsystems of the machine. With regard to the technical background, reference is now made to the figures, of which Figs. 1 to 3 show an example of a combustion machine based on combustion, in which embodiments of the present invention can be used. One of ordinary skill in the art will recognize that the present invention is not limited to this type of use. As mentioned, the present invention can be used in turbomachinery based on combustion, for example, in the machines used for power generation and in aircraft, in steam turbines and other types of rotating machinery. The examples given are not intended to limit the type of turbomachine.
    Fig. 1 is a schematic diagram of a combustion based turbomachine 10. Generally, combustion based turbomachinery operates by extracting energy from a compressed stream of hot gas produced by the combustion of fuel in a stream of compressed air. As illustrated in FIG. 1, the combustion-based turbomachine 10 may be comprised of an axial compressor 11 mechanically connected by a common shaft or rotor to a downstream turbine section or turbine 12 and one between the compressor 11 and the turbine 12 arranged combustion chamber 13 may be constructed.
    FIG. 2 shows a view of an example of a multi-stage axial compressor 11 that can be used in the combustion combustion engine of FIG. 1. As shown, the compressor 11 may include multiple stages. Each stage may include a series of compressor blades 14 followed by a series of compressor blades 15. Thus, a first stage may include a series of compressor blades 14 rotating about a central axis and followed by a series of compressor vanes that remain stationary during operation.
    FIG. 3 shows a partial view of an example of a turbine section or turbine 12 that may be used in the combustion machine of FIG. 1. The turbine 12 may include multiple stages. As examples, three stages are shown, but there may be more or fewer stages in the turbine 12. A first stage includes a plurality of turbine blades or turbine rotor blades 16 that rotate about the shaft during operation and a plurality of nozzles or turbine vanes 17 that remain stationary during operation. The turbine vanes 17 are generally spaced apart circumferentially and fixed about the axis of rotation. The turbine blades 16 may be mounted on a turbine wheel (not shown) so as to rotate about the shaft (not shown). A second stage 12 is also shown. The second stage includes a plurality of circumferentially spaced turbine nozzles 17, which are joined by a plurality of circumferentially spaced turbine blades 16, which are also rotatably mounted on a turbine wheel. A third stage is also shown and also includes a plurality of turbine nozzles in 17 and vanes 16. Note also that the turbine nozzle 17 and the turbine blades 16 are in the hot gas path of the turbine 12. The direction of the flow of hot gases through the hot gas path is indicated by the arrow. One of ordinary skill in the art will recognize that the turbine 12 may have more or, in some cases, fewer stages than shown in FIG. Each additional stage may include a series of turbine nozzles 17, followed by a series of turbine blades 16.
    In an example of operation, rotation of the compressor blades 14 within the axial compressor 11 may compress an airflow. In the combustion chamber 13, energy can be released when the compressed air is mixed with fuel and ignited. The resulting flow of hot gases from the combustor 13, which may be referred to as working fluid, is then directed via the turbine blades 16, the flow of the working fluid inducing rotation of the turbine blades 16 about the shaft. This converts the energy of the working fluid flow into the mechanical energy of the blades and, because of the connection between the blades and the shaft, the rotating shaft. The mechanical energy of the shaft can then be used to rotationally drive the compressor blades 14 so that the necessary supply of compressed air is generated and, for example, to a generator to generate electricity.
    Note that blades in the rear stages of gas turbines are extended to maximize the energy extracted from the hot gas stream. As the blades become longer, the blade 25 is twisted about the radial axis of the blade and may have other curved surface contours for improved fluid dynamic performance, as previously discussed. Another problem of larger turbine blades is the effect of such a relatively large mass due to rotation of the blade under an extremely high temperature. To counter this, the blade 25 can be tapered to reduce weight. However, active internal cooling of the blade 25 remains an important component in combating the extreme thermal and mechanical stresses experienced by these larger blades. Note that these stresses can cause creep problems and / or lead to the blade escaping due to deformation, which can adversely affect the fluid dynamic performance and blade life of the components.
    The fluid dynamics considerations associated with larger blades and the geometric constraints they exert make the production of certain conventional types of cooling structures difficult, expensive and / or impractical. Note that cooling of large turbine blades in the prior art is typically achieved by drilling radial holes in the blade 25 from the outside tip or foot. However, the highly warped and tapered profile associated with the larger blades makes drilling of long radial, straight lines impossible or too expensive. More specifically, conventionally formed straight or linear radial cooling channels can not be drilled from the platform to the tip so as to be able to supply certain areas of the blade 25 from the tip to the foot, due to the curved and tapered profile of the blade 25. The casting of such cooling channels is so expensive for several reasons that it is not worthwhile, among other things, because of the fact that core anchors, which are used in the casting of the cooling channels, are so easily damaged in the molds and consequently are incorrectly poured. This requires very targeted and efficient cooling solutions. The reduction in blade cross-sectional area available for drilling radial holes is a function of blade twist and taper. Greater blade twisting and tapering results in a smaller cross-sectional area available for drilling radial cooling holes. Cooling the large, heavily twisted and tapered blade by using a straight bore may not provide optimal cooling efficiency. What is especially missing is the cooling at the leading edge and the trailing edge of the blade. This prevents the use of such blades in a high firing temperature application as well as in a slow coolant flow design.
    Figures 4 to 8 show views of a turbine blade 16 of the type described above, in which embodiments of the present invention may be used. Note that these figures are presented in conjunction with FIGS. 9 and 10 to illustrate common geometrical limitations that bring internal cooling structures to their limit. As shown, the bucket 16 has a foot 21 by which the bucket 16 is attached to a rotor disk. The foot 21 may include a dovetail configured to be secured in a corresponding dovetail groove in the outer edge of the rotor disc. The foot 21 may further include a shank extending between the dovetail and a platform 24 disposed at the dovetail joint of the blade 25 and the foot 21 and defining a portion of the inner boundary of the flow path through the turbine 12. Note that the blade 25 is the active component of the blade 16, which catches the flow of working fluid and causes the rotor disc to rotate. Although the blade in this example is a blade 16, it is to be understood that the present invention may be applied to other types of blades within the fluid flow machine 10 unless otherwise specified, including the vanes 17, which may have similar structures. It can be seen that the blade 25 of the blade 16 has a concave pressure side surface 26 and a circumferentially or laterally opposite convex suction side surface 27 each extending axially between opposing upstream and downstream edges 28,29. The side surfaces 26 and 27 also extend in the radial direction from the platform 24 to an outer tip 31.
    The blade 25 may have a curved or contoured shape extending between the inboard end of the blade 25 (i.e., where the blade 25 extends radially from the platform 24) and the outboard tip 31. Note that certain regions or regions (or, as referred to herein, "target regions 41") within the blades of radially aligned cooling channels that extend along linear paths or reference lines can no longer be achieved when the curved or contoured shape of the Sheet 25 is more pronounced. Such target areas 41 exist because of the geometric constraints imposed by a modern aerodynamic design in both the outboard and inboard areas of the blade 25. These target areas 41 may also be close to the inflow and outflow edges, as discussed below of the sheet 25, especially near the trailing edge, because of the narrowing of the sheet profile, which is more pronounced near this end. As shown in Figs. 4 and 5, the blade 25 gradually tapers in its course from the platform 24 to the outer tip 31, which limits the use of linear cooling channels 33 in this area. The taper includes an axial taper that narrows the distance between the leading edge 28 and the trailing edge 29 of the blade 25, as shown in FIG. 4, as well as a circumferential taper, which is the thickness of the blade 25 between the suction side surface 26 and the pressure-side surface 27 is reduced. As shown in Figures 6 through 8, the contoured shape of the sheet 25 may further include a twist about a longitudinal axis of the sheet 25 (i.e., in the course of the sheet in the radial direction with respect to the turbine). The twist is typically designed to vary a stagger angle for the blade 25 generally between the inboard end and the outboard tip. However, it should be noted that the combined effects of the tapering and twisting structures imply a further limitation on the areas within the sheet that can be achieved using linearly formed cooling channels, such as the linear cooling channels 33 illustrated in FIG. 5 and FIGS as shown, can only be positioned within the central column of the blade 16. As also shown, the linear refrigerant passageway 33 may extend between a coolant inlet 35 extending through the root 21 of the bucket 16. At the other end of the linear refrigerant passageway 33, an outlet port 37 for discharging the refrigerant that moves through the blade 16 may be formed.
    For the purposes of the description, as further illustrated in FIG. 4, the blade 25 may be described as having a leading half and a tracked half defined on either side of an axial centerline 45. The axial centerline 45, as used herein, may be formed by connecting the centers 46 of the camber lines 47 of the sheet 25 (see FIG. 6). In addition, the blade 25 may be described as having two radially staggered portions defined within and outside a radial centerline 48 of the blade 25. As used herein, the inboard portion extends between the root and the radial centerline 48, while the outboard portion extends between the radial centerline 48 and the outboard tip 31.
    As shown in Figs. 9 and 10, the blade 25 may also include a radial diffraction component defining an arc or other curvature along the longitudinal axis of the blade 25. Note that FIGS. 9 and 10 also show blade structures that include distortion and taper that, when viewed in conjunction with the added curvature of the diffractive component, clearly demonstrate the inherent obstacles that arise when using linear cooling channels 32. to achieve all internal regions of the blade 25, which require active cooling during operation of the turbine or turbomachine.
    Figures 11 and 12 are side and perspective views of a turbine rotor blade 16 showing conventional linear cooling channels 33, while Figures 13 and 14 provide side and perspective views of a rotor disk 16 for comparison, the curved cooling channels 34 according to Embodiments show. As shown in Figs. 11 and 12, the area which can be achieved by linear drilling within a tapered and curved blade 25 is limited to the central parts of the blade 25. Thus, the target areas 41, which can not be reached by the linear cooling channels 33, are substantial, especially along both the leading and trailing edges. In contrast, Figs. 13 and 14 show a significantly enlarged portion of a cooling supply made possible by the curved cooling channels 34 of the present invention. As shown, the curved cooling channels 34 may extend into the highly curved and narrow regions adjacent the trailing edge of the blade 25. Further, these curved cooling channels 34 may be formed to extend below and near the surface of either the pressure or suction side surfaces. Specifically, the curved structure of the curved cooling channels 34 allows them to correspond to the curvature of each of the curved surface portions of the sheet 25 and to follow more closely. This may allow a much denser arrangement with respect to the numerous contoured surfaces of the sheet 25 than would otherwise be possible with a linear structure. Note that by allowing internal cooling circuits to be located closer to the surface, the cooling efficiency of the coolant is increased.
    Figures 15, 16 and 17 are perspective views of turbine blades 16 with curved cooling channels 34 according to several alternative embodiments of the present invention. As shown, the curved cooling passages 34 may be configured to extend between an upper end and a downstream end to have a predefined target area 41 therebetween (ie, a region inaccessible to linear cooling passages 33) is). According to preferred embodiments, several of the curved cooling channels 34 may be provided. Alternatively, one or more of the curved cooling channels 34 may be used in conjunction with one or more linear cooling channels 33. As shown in FIG. 15, according to one embodiment, the curved cooling channels 34 may extend in parallel from an inboard position where one or more of the curved cooling channels 34 forms a branch 39 from a central region of a linear segment or linear cooling channel 33. From mid-region connection 39, the curved cooling channels 34 may extend to an outboard location where they each open into an outlet port 37.
    As shown in Figs. 15 and 16, the curved cooling passages 34 may have an upstream upper end which is central, i.e., the upper end of the flow path. which is arranged near the axial center line. The upstream end of the curved cooling channels 34 may also be located near or at the inboard end of the blade 25, where it may be connected to a coolant inlet 35. The coolant inlet 35 may be configured as a hollow conduit which extends through the root 21 of the blade 16 so as to be connected to a coolant source. From the upstream end, the curved cooling channels 34 may extend in a generally radial direction while having a curvature that curves in accordance with the contoured shape of the blade 25. As shown, the curved cooling channels 34 may have a curvature that bends to the outlet openings 37 formed along the trailing edge 29 of the sheet 25. The outlet openings 37 may be radially spaced along the trailing edge 29 of the sheet 25. The upstream ends of the one or more curved cooling passages 34 may be disposed within the leading half / leading half of the blade 25, and the outlet ports 37 to which the curved cooling passages 34 are connected may be formed through an outer surface formed in the trailing half / outflow half of the sheet 25 is arranged. Note that such a structure provides a long convection path for the coolant, which affects its efficient use. As shown in Fig. 17, outlet openings 37 may also be formed in the outer tip of the sheet 25. It should also be noted that the discharge ports 37 may be formed on each outer surface of the sheet 25 as needed, including discharge ports formed as film cooling outlets formed on the pressure-side surface 26 or the suction-side surface 27 of the sheet 25. Although not shown, one or more outlet ports 37 may be formed on the leading edge 28 of the blade 25 so as to connect curved cooling passages to that region.
    According to a preferred embodiment, the curved cooling channels 34 may be designed as a feature formed after casting. As used herein, a feature formed after casting is a feature added to the sheet after a conventional molding process has formed the main body to the component. According to certain embodiments, the curved cooling channels 34 are formed using a steerable or controllable electrochemical removal process. According to other embodiments, the curved cooling channels 34 of the present invention may be formed using a controlled discharge electric discharge method and / or 3D printing methods. According to certain embodiments, the blade 16 in which the curved cooling channels 34 are used may be one designed to operate in a rear or downstream lower row of long blades, for example those in multi-stage turbines.
    One of ordinary skill in the art will recognize that many varying features and structures described above with respect to several embodiments may also be selectively applied to form the other possible embodiments of the present invention. For the sake of brevity and in the knowledge of one of ordinary skill in the art, not all possible steps have been specified or discussed in detail, but all combinations and possible embodiments encompassed by the several claims below or otherwise are intended to be part of the present invention. Moreover, those skilled in the art will be able to derive improvements, modifications and modifications from the above limitation of several embodiments of the invention. Such improvements, changes and modifications that are within the skill of the art are also intended to be covered by the appended claims. It is further to be understood that the above teachings are only to refer to the described embodiments of the present application and that numerous changes and modifications may be made thereto without departing from the spirit and scope of the application which are covered by the following claims and their equivalents.
    A blade for a turbine of a turbomachine, which includes a blade, wherein the blade has a leading edge, a trailing edge, an outer tip and an inner end. The blade may further include a cooling structure having a plurality of cooling channels for receiving and directing a coolant. The plurality of cooling channels may include a linear cooling channel and a curved cooling channel. The blade may also have a contoured shape defined by the blade between the inboard end and the outboard tip, wherein the contoured shape is configured to include a target region that extends from a linear datum line that extends radially from a location at the inboard end extends from the leaf, can not be reached. The curved cooling passages may be configured to extend between a downstream end and a downstream end so as to intersect the target area therebetween.
    权利要求:
    Claims (10)
    [1]
    A blade for a turbine of a gas turbine having a blade, the blade having a leading edge, a trailing edge, an outer tip and an inner end where the blade is attached to a foot adapted to engage the turbine blade with a blade Rotor disc, the sheet further comprising a cooling structure, which includes a plurality of elongated cooling channels for receiving and passing coolant through the sheet, wherein the blade further comprises:a contoured shape defined by the blade between the inboard end and the outboard tip, the contoured shape configured to have a target area that can not be reached by a linear datum line extending radially through the blade; anda curved cooling passage configured to extend between a downstream end and a downstream end so as to cross the target area therebetween;a coolant inlet configured to be in fluid communication with the lower end of the curved cooling channel.
    [2]
    2. A blade according to claim 1, wherein the coolant inlet passes through the root of the blade so that it is connected to a coolant source;wherein the blade comprises a blade and the curved cooling channel comprises a feature formed after casting; andwherein the linear reference line comprises one extending between the outer tip and the inner end of the blade.
    [3]
    A blade according to claim 1 or 2, wherein the contoured shape of the blade comprises at least one of the following:a concave-shaped pressure-side surface and a convex-shaped suction-side surface, which are connected to each other along the inflow and outflow edges;a radial diffraction component by which an arc is formed along a longitudinal axis of the sheet;twisting along a longitudinal axis of the blade, the twist being designed to gradually change a stagger angle for the blade between the inboard end and the outboard tip; anda taper along a longitudinal axis of the blade.
    [4]
    4. A blade according to claim 3, wherein the taper comprises at least one of the following:an axial taper gradually decreasing a distance between the leading edge and the trailing edge between the inboard end and the outboard tip of the sheet; anda taper in the circumferential direction by which a thickness between the pressure-side surface and the suction-side surface between the inner end and the outer tip of the sheet is gradually reduced.
    [5]
    A blade according to claim 3 or 4, wherein the structure of the linear unreachability of the target region comprises a combination of at least two of the following: twist about the longitudinal axis of the blade; Rejuvenation along the longitudinal axis of the leaf; radial diffraction component by which the arc is defined along a longitudinal axis of the sheet; andwherein a curvature of the curved cooling channel is shaped to correspond to a curvature of the contoured shape of the sheet.
    [6]
    A blade according to any one of the preceding claims, wherein the blade comprises a leading half and a trailing half defined on each side of an axial centerline connecting the centers of blade camber lines;the blade comprising radially staggered portions defined inwardly and outwardly from a radial centerline of the blade, an inboard portion extending between the root and the radial centerline, and an outward portion between the radial centerline and the outboard tip extends; andwherein the cooling structure comprises a plurality of the curved cooling channels.
    [7]
    The blade of claim 6, wherein the upstream upper end of each of the curved cooling channels comprises one located near the axial centerline and the inboard end of the blade; andwherein the curved cooling channels include a bend that bends to outlet openings formed on at least one of: the leading edge of the sheet; the trailing edge of the sheet.
    [8]
    A blade according to claim 6 or 7, wherein each of the curved cooling channels is parallel with respect to the other curved cooling channels and along a curve which bends towards the trailing edge of the blade;wherein the outlet openings have a radial distance along the trailing edge of the sheet.
    [9]
    A bucket according to any one of the preceding claims, wherein all curved cooling passages extend from an inboard position, where they are each connected at a mid-range connection to one of the linear cooling passages, parallel to an outboard position where they are respectively connected to one of the discharge ports are.
    [10]
    10. A blade according to one of the preceding claims, wherein the curved cooling channels have a formed after casting feature, which is formed after the casting of the blade by a removal method, andwherein the ablation process comprises a controllable electrochemical ablation process or a spark erosion-based ablation process.
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    同族专利:
    公开号 | 公开日
    CN205532726U|2016-08-31|
    DE102015121651A1|2016-06-30|
    JP2016125484A|2016-07-11|
    US20160186574A1|2016-06-30|
    引用文献:
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    US2687278A|1948-05-26|1954-08-24|Chrysler Corp|Article with passages|
    US3017159A|1956-11-23|1962-01-16|Curtiss Wright Corp|Hollow blade construction|
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    US20130052035A1|2011-08-24|2013-02-28|General Electric Company|Axially cooled airfoil|EP3028793A1|2014-12-04|2016-06-08|Siemens Aktiengesellschaft|Method for manufacturing a rotor blade|
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    2019-03-15| AZW| Rejection (application)|
    优先权:
    申请号 | 申请日 | 专利标题
    US14/583,865|US20160186574A1|2014-12-29|2014-12-29|Interior cooling channels in turbine blades|
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