Gas turbine nozzle having cooling holes and turbine

Abstract

A nozzle segment for a gas turbine includes an inner platform (23) and an outer platform (25) and a number of airfoils (27, 29) arranged between the inner platform (23) and the outer platform (25). A set of platform film cooling holes (37) is provided, including at least one platform film cooling hole on a surface of the inner platform or on a surface of the outer platform facing the hot gas flow path (31). The ratio of the tangential distance (Δγhole) of the platform film cooling hole (37) from the pressure side of the respective airfoil (29) to the width (Δγ) of the tangential hot gas flow path (31) at the platform film cooling hole (37) is comprised between 0 and 0.5.

Description

[Technical field]

[0001] FIELD OF THE DISCLOSURE The present disclosure relates to gas turbine engines. In particular, disclosed herein are turbine components that require film cooling, such as a stationary inlet gas nozzle located between a combustor and a high pressure turbine wheel. [Background technology]

[0002] In improving the thermal efficiency of a gas turbine engine, high combustion temperatures are desirable because increasing the hot part of the thermodynamic cycle increases thermal efficiency.

[0003] The high combustion gas temperatures require cooling of the turbine components closest to the combustion chamber, i.e., the combustor, typically the first stage nozzle through which the high pressure, hot combustion gases from the combustion chamber flow towards the first turbine wheel of the high pressure turbine. Film cooling of the components is typically used to prevent or reduce thermal damage and wear of the components through which the hot combustion gases flow. Film cooling is achieved by supplying cooling air through holes manufactured in the turbine components and having hole exits on the surface of the turbine component to be cooled. These holes are commonly referred to as "film cooling holes" or "film holes."

[0004] Typically, film cooling holes are manufactured in the airfoil features of a stationary nozzle located between the combustion chamber and the high pressure turbine wheel, and in some turbines, film cooling holes are also provided in concentrically arranged inner and outer platforms between which are arranged the airfoil features that define the gas flow path.

[0005] Cooling effectiveness may be affected by the location and orientation of the film cooling holes relative to the airfoil.

[0006] Improved cooling hole designs are needed to improve cooling efficiency to enable higher temperatures in the thermodynamic cycle and / or reduce thermal damage to turbine components. Summary of the Invention

[0007] Disclosed herein is a turbine component, and more specifically, a nozzle segment (nozzle sector) for a gas turbine including two or more airfoils defining one or more nozzles, the one or more nozzle segments arranged circumferentially about a turbine shaft to form a first set of stationary high pressure nozzles that direct a flow of hot pressurized combustion gases into a first turbine wheel.

[0008] In the embodiments disclosed herein, the nozzle segment includes an inner platform and an outer platform, and the airfoil is disposed between the inner platform and the outer platform.

[0009] The inner platform, outer platform and airfoil are manufactured as a single integral component by additive manufacturing.

[0010] Each airfoil includes a leading edge, a trailing edge, a pressure side, and a suction side. Each pair of back-to-back airfoils defines a nozzle, or hot gas flow path.

[0011] More specifically, each hot gas flow path is defined between an inner platform, an outer platform, the suction side of one airfoil, and the pressure side of another adjacent airfoil.

[0012] The nozzle segments also include film cooling holes, including at least one inner film cooling hole in each flowpath, located on either the inner platform surface or the outer platform surface, or both, facing the respective hot gas flowpath.

[0013] The film cooling holes are formed as the nozzle segments are produced by additive manufacturing, thereby overcoming the mechanical limitations of other manufacturing methods such as drilling or electrical discharge machining (EDM).

[0014] To improve cooling efficiency, the film cooling holes are shaped film cooling holes and are arranged such that the ratio of the tangential distance of the film cooling hole from the pressure side of each airfoil to the width of the tangential hot gas flow path at the film cooling hole is included between 0 and 0.5. Further, the stagger angle of each airfoil is less than or equal to 85°. Additionally, the ratio of the vane flow path width to the axial chord of each such flow path is included between 0 and 0.5.

[0015] A "shaped film cooling hole", also referred to as a "diffusion shaped hole" or "diffusion cooling hole", is a hole that has a diverging shape at the exit end that is neither circular nor elliptical and aids in spreading the cooling flow. Shaped film cooling holes improve cooling efficiency by providing greater surface coverage of the component to be cooled. A typical shaped film cooling hole may consist of a cylindrical first channel portion and have a cross section that increases in the upstream-to-downstream direction near the exit end.

[0016] Cooling holes that satisfy the geometric conditions outlined above have been found to be surprisingly effective in improving the cooling efficiency of an airfoil. Indeed, due to pressure gradients between the pressure and suction sides of adjacent airfoils, the cooling medium flowing from the cooling holes in the platform surface tends to move away from the pressure side and towards the suction side of the opposing airfoil. This results in insufficient thermal protection of the pressure side of the airfoil. By locating the film cooling holes near the pressure side of the airfoil, the cooling efficiency of the pressure side of the airfoil is improved. This positioning, coupled with the use of shaped film cooling holes, provides a synergistic effect in that the diffusion effect of the shaped film cooling holes is applied adjacent the pressure side of the airfoil. This location and diffusion effect promotes the cooling effect of the pressure side of the airfoil, despite the pressure gradient of the gas flowing between adjacent airfoils that tends to transport the cooling air away from the pressure side of the adjacent airfoil and towards the suction side.

[0017] In addition to the shaped film cooling holes that satisfy the geometric conditions described above, additional film cooling holes that do not satisfy the conditions described above may be provided in one or more of the flow passages, i.e., closer to the centerline of the flow passage and / or even further left on the inner and / or outer platform surfaces when looking aft from the front of the nozzle segment.

[0018] Further features and embodiments are described hereinafter with reference to the accompanying drawings and set forth in the appended claims. [Brief description of the drawings]

[0019] Reference will now be made briefly to the accompanying drawings, in which: [Figure 1] FIG. 1 is a schematic diagram of an aeroderivative gas turbine engine in which turbine components according to the present disclosure may be used. [Diagram 2] FIG. 2 illustrates an axonometric view of a nozzle segment according to the present disclosure. [Diagram 3] FIG. 3 illustrates an axonometric view of a nozzle segment according to the present disclosure. [Figure 4] FIG. 4 illustrates an axonometric view of a nozzle segment according to the present disclosure. [Diagram 5] FIG. 5 illustrates an axonometric view of a nozzle segment according to the present disclosure. [Figure 6] FIG. 6 is a schematic cross-sectional view of a hot gas flow path formed within the nozzle segment of FIGS. [Figure 7] FIG. 7 is a schematic representation of a cross-section of the inner platform and one of the outer platforms along the axis of a film cooling duct. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0020] FIG. 1 illustrates a schematic diagram of an exemplary gas turbine engine 1 in which a nozzle segment according to the present disclosure may form a fixed nozzle between a combustor and a high pressure turbine wheel.

[0021] 1 is an aeroderivative gas turbine engine including a compressor section 2, a combustor 3, and a turbine section 4. Gas turbine engine 1 is illustrated as one example of a gas turbine engine in which turbine components according to the present disclosure may be used. However, those skilled in the art of turbomachinery will appreciate that the advantageous features of the nozzle segments according to the present disclosure may be used in gas turbine engines of different constructions and configurations, including, for example, different numbers of turbine wheels, compressors, and shafts.

[0022] In the exemplary embodiment of FIG. 1, the compressor section 2 includes a low pressure compressor 2A and a high pressure compressor 2B. In the exemplary embodiment of FIG. 1, the turbine section 4 includes a high pressure turbine 4A, a low pressure turbine 4B, and a power turbine 4C. The high pressure turbine 4A is drivingly connected to the high pressure compressor 2B through a first shaft 5 such that power generated by the high pressure turbine 4A drives the high pressure compressor 2B. The low pressure turbine 4B is drivingly connected to the low pressure compressor 2A through a second shaft 6 such that power generated by the low pressure turbine 4B drives the low pressure compressor 2A. A third shaft 7 may drivingly connect the power turbine 4C to a load, which may include a gas compressor or compressor train, a generator, or any other type of driven machine.

[0023] Compressed air delivered by compressor section 2 to combustor 3 is mixed with fuel and the fuel-air mixture is combusted to generate a flow of hot pressurized combustion gases that are expanded sequentially in high pressure turbine 4A, low pressure turbine 4B, and power turbine 4C to generate power for driving compressor section 2 and load 8.

[0024] The flow of hot pressurized combustion gases from combustor 3 is delivered to the high pressure turbine 4A through an annular set of stationary airfoils which, together with inner and outer platforms, form a nozzle, shown diagrammatically at 4D, which directs the combustion gases towards the turbine wheel.

[0025] The annular arrangement of nozzles may be formed by a series of nozzle segments arranged annularly about the axis AA of the gas turbine engine 1. Each nozzle segment may include inner and outer platforms and a number of airfoils, e.g., first and second airfoils, disposed therebetween. References to "first" and "second" airfoils are conventional and are intended merely to distinguish one airfoil from the other and should not be construed as limiting the present disclosure in any manner.

[0026] 2, 3, 4, 5 and 6 show in detail one of the nozzle segments 11, which when assembled form an annular arrangement of the gas nozzle 4D. In the embodiment illustrated herein, each nozzle segment includes two airfoils, but other embodiments may include three or more airfoils in each nozzle segment. More specifically, FIGS. 2 and 3 show axonometric views from the inlet and outlet sides, respectively, of a nozzle segment with an inner platform at the bottom of the view. FIGS. 4 and 5 show axonometric views from the inlet and outlet sides, respectively, of the same nozzle segment with an outer platform at the bottom of the view.

[0027] More specifically, the nozzle segments 11 include an inner platform, or inner platform portion 23, and an outer platform, or outer platform portion 25. When a plurality of nozzle segments 11 are assembled together in an annular arrangement surrounding the turbine axis AA, the inner platforms, or inner platform portions 23 of the nozzle segments 11 form a complete annular inner platform of the gas inlet nozzle of the high pressure turbine. Similarly, the outer platforms, or outer platform portions 25 of the nozzle segments 11 form a complete annular outer platform of the gas inlet nozzle of the high pressure turbine.

[0028] In the following, elements 23 and 25 will be referred to simply as the inner platform and the outer platform, respectively, even though they may represent only a portion of the complete inner platform and outer platform.

[0029] The first airfoil 27 and the second airfoil 29 are disposed between the inner platform 23 and the outer platform 25 and extend from a surface 23A of the inner platform 23 facing the outer platform 25 to a surface 25A of the outer platform 25 facing the inner platform 23. As seen in FIG. 6, the first airfoil 27 has a suction side 27S and a pressure side 27P. Specifically, the suction side 27S is convex and the pressure side 27P is concave. The suction side 27S and the pressure side 27P blend together to form a leading edge 27L and a trailing edge 27T of the first airfoil 27. Similarly, the second airfoil 29 has a convex suction side 29S and a concave pressure side 29P. The suction side 29S and the pressure side 29P blend together at a leading edge 29L and a trailing edge 29T of the second airfoil 29.

[0030] The hot gas flow path 31 extends between the suction side 27S of the first airfoil 27, the pressure side 29P of the second airfoil 29, and the facing surfaces 23A, 25A of the inner and outer platforms 23, 25.

[0031] As can be seen in this figure, the first airfoil portion 23 and the second airfoil portion 29 are preferably equal to one another.

[0032] Cooling air flowing through suitable channels within the first and second airfoils 27, 29 may be delivered to conventional film cooling holes (not shown) on the airfoil surfaces.

[0033] In addition to conventional cooling holes on the airfoils 27 , 29 , the nozzle segment 11 includes conventional film cooling holes 35 on the inner and outer platforms 23 , 25 located near the entrance and exit of the hot gas flow path 31 .

[0034] In accordance with the present disclosure, at least an additional set of film cooling holes including at least one platform film cooling hole 37 is provided in either or both of the inner platform 23 and the outer platform 25 near a location near the pressure side 29P of the second airfoil 29, i.e., near a concave surface of the second airfoil profile 29 facing the interior of the hot gas path 31. In the illustrated embodiment, three additional platform film cooling holes 37 are provided at three locations along the direction of the hot gas path 31. In other embodiments, at least one platform film cooling hole may be provided only in the inner platform 23 or only in the outer platform 25.

[0035] The platform film cooling holes 37 are shown in detail in FIG. 6, which shows a cross-section of the nozzle segment 11 along a surface extending between and parallel to the inner platform 23 and the outer platform 25. The features of the nozzle segment 11 referred to below and referring to the platform film cooling holes 37 can be replicated on both the inner platform 23 and the outer platform 25, so FIG. 6 can show either the inner platform and the outer platform surface 23A, 25A, or the other. For simplicity and convenience, FIG. 6 below will represent the surface 23A of the inner platform 23. However, it should be understood that the parameters defined below and the location of the platform film cooling holes described below can be referenced in the same way to the outer platform. In other words, the outer platform 25 can be provided with an arrangement of platform film cooling holes 37 similar to the arrangement of the platform film cooling holes 37 provided on the inner platform 23. However, it should be understood that the location of the platform film cooling holes 37 on the inner platform 23 may be different from the location of the platform film cooling holes 37 on the outer platform 25, provided that the geometric constraints described herein below are met.

[0036] In some embodiments, each platform film cooling hole 37 may be a shaped hole, as seen in Figure 6. As understood herein, each platform film cooling hole 37 is the exit end of a film cooling duct that penetrates the thickness of the respective inner platform 23 or outer platform 25. Figure 7 illustrates a schematic cross-sectional view of one of the inner and outer platforms 23, 25 along the axis of a film cooling duct 38 that terminates in a platform film cooling hole 37.

[0037] The location of each platform film cooling hole 37 is selected to provide maximum cooling performance. Before defining the location of each platform film cooling hole 37, some parameters related to the nozzle segment 11 and the platform film cooling holes 37 are defined, and with particular reference to Figure 6, the dashed dotted line AA indicates the axial direction, i.e., the direction parallel to the axis AA of the gas turbine engine 1 (see also Figure 1). For ease of viewing the drawings, the geometric parameters described below and defining the location of the platform film cooling holes 37 are shown in Figure 6 only for the mid-platform film cooling holes 37.

[0038] Reference is now made to definitions of the geometric parameters used below to define the shape and location of the platform film cooling holes 37, where the "axial direction" represented by dashed dotted line AA is the direction parallel to the axis AA of the gas turbine engine 1 when the nozzle segment 11 is mounted on the gas turbine engine 1. The tangential direction illustrated by dashed dotted line TT is the direction perpendicular to the axial direction AA. Cax is the distance between the trailing edge 27T, 29T and the leading edge 27L, 29L of each airfoil 27, 29 in the axial direction AA, and is referred to as the "axial chord."

[0039] Cax,hole is the distance in the axial direction AA of the platform film cooling hole 37 from the leading edges 27L and 29L of the first and second airfoils 27 and 29.

[0040] Reference symbol St denotes the vane passage width, which is also referred to as the width of the vane passage, i.e., the throat or opening of the hot gas passage 31 at the trailing edge 29T of the second airfoil 29. That is, the vane passage width is the distance from the suction side 27S of the first airfoil 27 to the trailing edge 29T of the second airfoil 29.

[0041] Δy is the tangential distance between the suction side 27S of the first airfoil 27 and the pressure side 29P of the second airfoil 29 along a line (perpendicular to the axial direction AA, i.e., tangential) passing through the center of the platform film cooling hole 37.

[0042] Δyhole is the tangential distance between the center of the platform film cooling hole 37 and the pressure side 29P of the second airfoil section 29.

[0043] The angle designated α, ie the angle between the chord CH and the tangent direction TT of the airfoil profiles 27, 29, is referred to herein as the staggered complement angle.

[0044] Line Ahole is the projection on the surface of the respective inner platform 23 or outer platform 25 of axis A (FIG. 7) of cooling duct 38 terminating in platform film cooling hole 37. Angle β, also called and referred to as the hole compound angle herein, is the angle between line Ahole and tangent direction TT.

[0045] According to one aspect of the present disclosure, the first airfoil 27 and the second airfoil 29 are shaped and positioned such that the angle α (i.e. the inclination of the chord of the airfoils 27, 29 with respect to the tangent direction TT) is comprised between 0° and 85°. In a preferred embodiment, such inclination is comprised between 0° and 80°, more preferably between 0° and 75°.

[0046] Furthermore, the ratio between the width St of the blade passage and the axial chord Cax is included in the range of 0 to 0.5, i.e.

[0047]

number

[0048] In a preferred embodiment, the aforementioned ratio is less than or equal to 0.45, more preferably less than or equal to 0.4.

[0049] In the nozzle segment 11 characterized by the geometrical features of the airfoils 27, 29 defined above, the platform film cooling holes 37, more precisely each platform film cooling hole 37, are arranged adjacent to the pressure side 29P of the second airfoil 29, i.e. at a distance Δyhole in the tangential direction TT therefrom, such that the ratio between said distance Δyhole and the distance Δy between the pressure side 29P of the second airfoil 29 in the tangential direction at the location of the platform film cooling hole 37 and the suction side 27S of the first airfoil 27 is between 0 and 0.5, preferably between 0 and 0.45, more preferably between 0 and 0.4, i.e.

[0050]

number

[0051] In a preferred embodiment, the, or each, platform film cooling hole 37 is preferably located deep within the hot gas flow path 31, i.e. at a relatively long distance from the leading edge 27L, 29L and from the trailing edge 27T, 29T. More specifically, the ratio between the distance Cax,hole of the platform film cooling hole 37 and the axial chord Cax is preferably between 0.15 and 0.95, i.e.

[0052]

number

[0053]

number

[0054]

number

[0055] The inclination δ (see FIG. 7) of the axis A of the duct 38 facing the surface 23A or 25A in the platform film cooling hole 37 is preferably comprised between 0° and 60°, preferably between 0° and 55°, more preferably between 0° and 50°. The hole combination angle β is preferably such that the difference between the hole combination angle and the stagger α is less than −15° or more than 25°, more preferably less than −7.5° or more than 27.5°, even more preferably less than −5° or more than 30°.

[0056] The geometric relationships outlined above provide particularly efficient film cooling of the nozzle segment 11 by the air-cooled film since the region of the vane passage 31 is located away from the leading and trailing edges and near the pressure side of the second airfoil portion 29.

[0057] If the nozzle segment 11 is manufactured by additive manufacturing, the above geometrical conditions can be advantageously met, since this manufacturing technique is not subject to the constraints and limitations of currently used techniques, such as, inter alia, electrical discharge machining.

[0058] Exemplary embodiments are disclosed above and illustrated in the accompanying drawings. It will be understood by those skilled in the art that various modifications, omissions, and additions may be made to what is specifically disclosed herein without departing from the scope of the invention as defined in the following claims.

Claims

1. 1. A nozzle segment for a gas turbine, comprising: an inner platform (23) and an outer platform (25); a plurality of airfoils (27, 29) disposed between the inner platform (23) and the outer platform (25), wherein the inner platform (23), the outer platform (25), and the plurality of airfoils (27, 29) are manufactured by additive manufacturing, each airfoil (27, 29) having a leading edge (27L, 29L), a trailing edge (27T, 29T), a pressure side (27P, 29P), and a suction side (27S, 29S), and a hot gas flow path (31) is formed between the inner platform (23), the outer platform (25), and each pair of successively arranged airfoils (27, 29) of the plurality of airfoils (27, 29); a set of platform film cooling holes (37) including at least one platform film cooling hole on a surface of the inner platform or on a surface of the outer platform facing the hot gas path (31); the cooling hole is a shaped cooling hole having a non-circular and non-elliptical diverging shape; a ratio of the tangential distance (Δyhole) of the platform film cooling hole (37) from the pressure side (29P) of each airfoil (29) to the width (Δy) of the hot gas flow path (31) tangentially at the platform film cooling hole (37) is comprised between 0 and 0.5; a stagger angle (α) of each of said airfoils is less than or equal to 85°; A nozzle segment, wherein the ratio of vane passage width (St) to axial chord (Cax) of each said hot gas passage (31) is comprised between 0 and 0.

5.

2. 2. A nozzle segment according to claim 1, wherein the ratio between the axial distance (Cax,hole) of said platform film cooling holes (37) from said leading edge (27L, 29L) and said axial chord (Cax) is comprised between 0.15 and 0.95, preferably between 0.2 and 0.925, more preferably between 0.25 and 0.

9.

3. 3. A nozzle segment according to claim 1 or 2, wherein the platform film cooling holes (37) have an axis (A) which forms an angle (δ) of 60° or less with a surface of the respective platform (23, 25) facing the hot gas path (31), preferably 55° or less, more preferably 50° or less.

4. 4. A nozzle segment according to any one of claims 1 to 3, wherein a difference between a compound angle (β) of the platform film cooling holes (37) and the staggered complement angle (α) of the airfoil is less than or equal to 15° or more than or equal to 25°, preferably less than or equal to -7.5° or more than or equal to 27.5°, more preferably less than or equal to -5° or more than or equal to 30°.

5. 5. A nozzle segment according to claim 1, wherein the ratio of the tangential distance (Δyhole) of the platform film cooling hole (37) from the pressure side of the respective airfoil (29) to the width (Δy) of the hot gas flow path (31) tangentially at the platform film cooling hole (37) is comprised between 0 and 0.45, preferably between 0 and 0.

40.

6. Nozzle segment according to any one of the preceding claims, wherein the staggered angle (α) is less than or equal to 80°, preferably less than or equal to 75°.

7. 7. A nozzle segment according to any one of claims 1 to 6, wherein the ratio between the vane passage width (St) and the axial chord (Cax) is comprised between 0 and 0.45, preferably between 0 and 0.

40.

8. 8. The nozzle segment of claim 1, wherein the set of platform film cooling holes (37) includes at least one platform film cooling hole on the inner platform (23) and at least one platform film cooling hole on the outer platform (25).

9. 9. The nozzle segment of claim 1, wherein the set of platform film cooling holes (37) comprises a plurality of platform film cooling holes on the inner platform (23) arranged in series along each hot gas flow path (31).

10. 10. The nozzle segment of claim 1, wherein the set of platform film cooling holes (37) comprises a plurality of platform film cooling holes on the outer platform (25) arranged in series along each hot gas flow path (31).

11. A gas turbine engine comprising at least one nozzle segment according to any one of the preceding claims.

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