Tip ski lift configuration

Abstract

To provide a method for a tip cap configuration allowing a cooling fluid flow out of a tip cap having a squealer tip.SOLUTION: The present embodiments set forth a blade including an airfoil 12 including: an outer tip 28 having a floor, a leading edge 51 and a trailing edge 50; a concave pressure sidewall 20 and a convex suction sidewall 22 that extend axially between corresponding leading and trailing edges and radially between the floor and a tip cap 36'. The airfoil further includes a tip cap extending from the floor of the outer tip 28 coextensively with the pressure and suction sidewalls and around the leading and trailing edges. The tip cap includes a squealer tip configuration including a suction side tip cap portion and a pressure side tip cap portion. The suction side tip cap portion and pressure side tip cap portion extend unequal distances above the floor, providing a cooling fluid flow out of the tip cap.SELECTED DRAWING: Figure 3

Description

Technical Field

[0001] This embodiment generally relates to an apparatus, method, and / or system for a turbine rotor blade. More particularly, but not limited to, this application relates to an apparatus and assembly for a turbine rotor blade having a ski tip configuration.

Summary of the Invention

[0002] This embodiment describes a blade comprising an airfoil having an outer tip with a floor, a leading edge, and a trailing edge, and concave pressure sidewalls and convex suction sidewalls that extend axially between the corresponding leading and trailing edges and radially between the floor and a tip cap. The airfoil further comprises a tip cap that extends around the leading and trailing edges with the same extent as the pressure sidewalls and the suction sidewalls from the floor of the outer tip. The tip cap comprises a ski tip configuration for reducing tip leakage and downstream mixing losses, and the ski tip configuration comprises a suction side tip cap portion and a pressure side tip cap portion. The pressure side tip cap portion comprises a step-down section, and the flow of cooling fluid passes over the pressure side tip cap portion of the tip cap due to a pressure gradient.

[0003] Another aspect of the embodiment describes a turbine engine comprising a blade. The blade comprises an airfoil having an outer tip with a floor, a leading edge, and a trailing edge, and concave pressure sidewalls and convex suction sidewalls that extend axially between the corresponding leading and trailing edges and radially between the floor and a tip cap. The airfoil further comprises a tip cap that extends around the leading and trailing edges with the same extent as the pressure sidewalls and the suction sidewalls from the floor of the outer tip. The tip cap comprises a ski tip configuration for reducing tip leakage and downstream mixing losses, and the ski tip configuration comprises a suction side tip cap portion and a pressure side tip cap portion. The pressure side tip cap portion comprises a step-down section, and the flow of cooling fluid passes over the pressure side tip cap portion of the tip cap due to a pressure gradient.

[0004] These and other features of the present application will become apparent upon consideration of the following detailed description of the preferred embodiments in conjunction with the drawings and the appended claims.

[0005] Exemplary aspects of the present disclosure have been developed to solve the problems described herein and / or other problems not discussed.

[0006] These and other features of the present embodiment will be more fully understood and appreciated by carefully considering the following more detailed description of the exemplary embodiments in conjunction with the following accompanying drawings.

Brief Description of the Drawings

[0007]

Figure 1

Figure 2

Figure 3

Figure 4

Figure 5

Mode for Carrying Out the Invention

[0008] It should be noted that the drawings of the present disclosure are not necessarily drawn to scale. The drawings are intended to show only typical aspects of the present disclosure and should not be regarded as limiting the scope of the present disclosure. In the drawings, like reference numerals between the drawings represent like elements.

[0009] The following examples of this embodiment may be described in relation to a particular type of turbine engine, but it will be understood by those skilled in the art that this embodiment is not limited to such usage and is applicable to other types of turbine engines, unless otherwise specifically limited. Further, it will be understood that specific terms may be used to refer to specific mechanical components within the gas turbine engine when describing this embodiment.

[0010] To the extent possible, common industrial technical terms are used and utilized in the same sense as their accepted meaning. However, such terms should not be construed narrowly, as it will be understood by those skilled in the art that specific mechanical components may often be referred to using different terms. Additionally, what may be described herein as a single component may be referred to in another context as consisting of multiple components, or what may be described herein as including multiple components may be referred to elsewhere as a single one. Therefore, when understanding the scope of this embodiment, pay attention not only to specific terms but also to the accompanying description, context, and in particular, the structure, composition, function, and / or usage of the components that may be provided in the appended claims.

[0011] In this specification, several descriptive terms may be used repeatedly, and it may be useful to define these terms at the beginning of this section. Therefore, these terms and their definitions are as follows, unless otherwise specified. As used herein, "downstream" and "upstream" refer to the direction of fluid flow, such as the flow of working fluid through the compressor section, combustor section, and turbine section of a gas turbine, or the flow of coolant through one of the component systems of an engine. The term "downstream" corresponds to the direction of fluid flow, and the term "upstream side" refers to the direction opposite to or against the direction of fluid flow. The terms "front" and "rear" refer to the direction with respect to the orientation of the gas turbine, unless otherwise specified, where "front" refers to the front of the engine or the compressor end, and "rear" refers to the rear of the engine or the turbine end. Additionally, considering the configuration of a gas turbine engine around the central axis, as well as this same type of configuration in several component systems, terms for describing the position relative to the axis are likely to be used. In this context, it will be understood that the term "radial direction" indicates movement or position perpendicular to the axis. In connection with this, it may be necessary to describe the relative distance from the central axis. In this case, for example, if a first component is located closer to the central axis than a second component, then in this specification, it is stated that the first component is "radially inside" or "inner" of the second component. On the other hand, if the first component is farther from the axis than the second component, then it can be stated in this specification that the first component is "radially outside" or "outer" of the second component. Further, it will be understood that the term "axial direction" refers to movement or position parallel to the axis. And finally, the term "circumferential direction" refers to movement or position around the axis.

[0012] Referring to the drawings, FIG. 1 is a schematic diagram of an exemplary turbomachine 90 in the form of a combustion turbine or gas turbine (GT) system 100 (hereinafter referred to as "GT system 100"). The GT system 100 includes a compressor 102 and a combustor 104. The combustor 104 includes a combustion region 105 and a fuel nozzle assembly 106. The GT system 100 also includes a turbine 108 and a compressor / turbine common shaft 110 (hereinafter referred to as "rotor 110"). In one embodiment, the GT system 100 is a 7HA.03 engine commercially available from General Electric Company of Boston, Massachusetts. The present disclosure is not limited to any one particular GT system and can be incorporated in connection with, for example, other HA, F, B, LM, GT, TM, and E class engine models of General Electric Company, as well as other engines including engine models of other companies. Further, the teachings of the present disclosure are not necessarily applicable only to GT systems and can be applied to other types of turbomachines, such as steam turbines, jet engines, compressors, and the like.

[0013] FIG. 2 is a cross-sectional view of an exemplary portion of a turbine 108 having four exemplary and non-limiting stages L0-L3, called L0-L3 for purposes of illustration only and not intended to convey any structure or particular machine, that can be used in the GT system 100 of FIG. 1. The four stages are called L0, L1, L2, and L3. Stage L0 is the first stage and the (radially) smallest of the four stages. Stage L1 is the second stage and the next stage axially. Stage L2 is the third stage and the next stage axially. Stage L3 is the fourth and last stage and the (radially) largest stage. The four stages are shown as an example only, and it should be understood that each turbine may have more or fewer than four stages. A set of stationary vanes or blades 112 cooperate with a set of rotating blades 114 to form each stage L0-L3 of the turbine 108 and define a portion of the flow path through the turbine 108. Each set of rotating blades 114 is coupled to respective rotor wheels 116 that circumferentially couple the rotating blades 114 to the rotor 110. That is, a plurality of rotating blades 114 are mechanically coupled to each rotor wheel 116 with circumferential spacing therebetween. The stationary blade section 115 comprises a plurality of stationary blades 112 circumferentially spaced around the rotor 110. Each blade 112 can comprise at least one end wall (or platform) 120, 122 connected to the airfoil 12. In the example shown, the blade 112 comprises a radially outer end wall 120 and a radially inner end wall 122. The radially outer end wall 120 couples the blade(s) 112 to the casing 124 of the turbine 108.

[0014] Another aspect of the present disclosure provides a blade for the final stage of a turbomachine as an embodiment of the present specification. The flow path on such a machine in the final stage is substantially conical, and the structures and features herein function similarly in the final stage blades having a conical flow path, as in other blade flow paths.

[0015] During operation, air flows through compressor 102, and compressed air is supplied to combustor 104. Specifically, the compressed air is supplied to fuel nozzle assembly 106 integrated with combustor 104. Fuel nozzle assembly 106 is in fluid communication with combustion region 105. Fuel nozzle assembly 106 is also in fluid communication with a fuel source (not shown in FIG. 1) and guides fuel and air into combustion region 105. Combustor 104 ignites and burns the fuel. Combustor 104 is in fluid communication with turbine 108 where the thermal energy of the gas flow is converted into mechanical rotational energy. Turbine 108 is rotatably coupled to rotor 110 to drive rotor 110. Compressor 102 is also rotatably coupled to rotor 110. In an exemplary embodiment, there are multiple combustors 104 and fuel nozzle assemblies 106. In the following description, unless otherwise specified, only one of each component will be described. At least one end of turbine 108 may extend axially away from rotating shaft 110 and may be attached to a load or machine (not shown), such as a generator and / or another turbine, without limitation.

[0016] Referring to FIG. 3, blade 114 includes airfoil 12. Airfoil 12 includes a generally concave pressure sidewall 20 extending between opposing leading edge 51 and trailing edge 50, and an opposing generally convex suction sidewall 22. Sidewalls 20 and 22 also extend radially between a root portion (not shown for ease of understanding of the embodiment) and outer blade tip 28. Sidewalls 20 and 22 are spaced apart over substantially the entire span of airfoil 12 to permit the flow of cooling fluid as described below.

[0017] The airfoil portion 12 can have an internal cooling fluid configuration. The internal cooling fluid configuration can include, for example, at least one internal flow path 30 (shown by dashed lines in FIG. 3), such as a serpentine flow path, for guiding cooling fluid air through the airfoil portion 12. Inside each internal cooling fluid flow path 30, turbulators can be provided that are formed to improve the effectiveness of the cooling fluid (not shown for ease of understanding of the embodiment). The cooling fluid from each internal cooling fluid flow path 30 can be discharged through a corresponding number of cooling fluid holes 34 (FIG. 3) in the floor 128 of the outer blade tip 28 as described below.

[0018] As shown in FIG. 3, the airfoil portion 12 includes an outer blade tip 28. The outer blade tip 28 includes a blade tip cap 36 that can be integrally formed on the radially outer ends of the positive pressure side wall 20 and the negative pressure side wall 22. The blade tip cap 36 surrounds an internal flow path or cavity 29. As shown, the blade tip cap 36 surrounds the outer blade tip 28 extending from the trailing edge 50 to the leading edge 51.

[0019] At the trailing edge 50, an opening 52 is formed in the blade tip cap 36 in the positive pressure side wall 20 proximate to the trailing edge 50. The opening 52 allows the cooling fluid to exit the blade tip cap 36. The opening 52 extends from the trailing edge 50 towards the leading edge 51 in the positive pressure side blade tip cap portion 137. The size of the opening 52 can be varied by any suitable distance along the positive pressure side blade tip cap portion 137 to allow the cooling fluid to flow out of the cavity 29 along with the tip leakage flow that is entrained into the ski blade tip configuration by operation. The flow of the cooling fluid is caused by the pressure gradient that essentially occurs from the positive pressure side to the negative pressure side of the blade during operation of the turbine blade and the airfoil portion.

[0020] The wingtip cap 36 includes a negative-pressure ski wingtip configuration and a positive-pressure ski wingtip configuration. The negative-pressure ski wingtip configuration and the positive-pressure ski wingtip configuration are formed by a negative-pressure side wingtip cap portion 136 and a positive-pressure side wingtip cap portion 137 having the same spread as the positive-pressure sidewall and the negative-pressure sidewall 20, 22. The negative-pressure side wingtip cap portion 136 is provided at a height Y from the floor 128 of the wingtip cap 28, and the positive-pressure side wingtip cap portion 137 is provided at a height X from the floor 128 of the outer wingtip 28 (see FIGS. 4 and 5). According to an embodiment of the present disclosure, height Y > height X. The difference in height reduces the rubbing of the blade wingtip with an adjacent casing (not shown). Also, the flow of the cooling fluid in the cavity 29 at the outer wingtip 28 can exit the cavity 29 through the opening 52 at the trailing edge 50 regardless of the source, and can also exit the cavity 29 over the positive-pressure side wingtip cap portion 137. See arrow F in FIG. 3.

[0021] The difference between height Y and height X is generally less than about 100 mils. In some embodiments of the present disclosure, the difference between height Y and height X is generally less than about 100 mils, or the difference between height Y and height X is generally less than about 90 mils, or even the difference between height Y and height X can generally be less than about 80 mils. Alternatively, the difference between height Y and height X can generally be less than about 70 mils, and in some aspects, the difference between height Y and height X can generally be less than about 60 mils. In other aspects of the present disclosure, the difference between height Y and height X is generally less than about 50 mils. Further, according to another embodiment of the present disclosure, the difference between Y and X can be in the range of about 10 mils to about 50 mils, or extends from about 10 to about 50 mils, still allowing the flow of the cooling fluid from the cavity 29 through the opening 52 and over the positive-pressure side wingtip cap portion 137 from the cavity 29 (see arrow F (FIG. 3)).

[0022] As shown in FIGS. 3, 4, and 5, the leading edge 51 region of the wing tip cap 36 defines a structure for height variation of the wing tip cap portions 136 and 137 to generate a difference between heights X and Y for generating a height change Z. The offset 37 provides a structure for the height change Z. The offset 37 is shown with the offset 37 in the positive pressure side wing tip cap portion 137 as a height step between the wing tip cap portions 136 and 137 in the leading edge 51 region. Depending on the height change Z, the positive pressure side wing tip cap portion 137 becomes lower than the negative pressure side wing tip cap portion 136, and a step-down section 237 is defined within the positive pressure side wing tip cap portion 137. The step-down section 237 extends from the offset 37 to the opening 52.

[0023] The configuration of the offset 37 can provide physical integrity to the blade 114 and the airfoil 12. The configuration of the offset 37 can also physically extend along the positive pressure side wing tip cap portion 137 any suitable distance to allow flow over the positive pressure side wing tip cap portion 137 and flow out of the opening 52. The step-down section 237 can extend across any portion of the positive pressure side wing tip cap portion 137. Thus, the length of the step-down section 237 can vary its length from the offset 37 to the opening 52. In one aspect of the embodiment, as in FIG. 3, the offset 37 is close to the leading edge or at the leading edge 51, and thus the step-down section 237 extends substantially over the positive pressure side wing tip cap portion 137. The further the offset 37 is from the leading edge 51 (conversely, the closer the offset 37 is to the trailing edge 50), the shorter the step-down section 237 becomes. The short step-down section 237 allows the flow of cooling fluid out of the cavity 29 over the positive pressure side wing tip cap portion 137, while the longer step-down section 237 provides more area for the flow of cooling fluid over the positive pressure side wing tip cap portion 137, which of course results in enhanced cooling fluid.

[0024] Offset 37 can take any suitable configuration that provides a change / difference in height step between the wingtip cap portion 136 and the wingtip cap portion 137. Offset 37 is shown in FIG. 5 with orthogonal steps (solid lines), inclined steps (dotted lines), and curves (dashed lines), all of which are within the scope of the embodiments and are used individually or in combination.

[0025] FIG. 5 shows a side view of the airfoil 12 having an offset 37 step configuration proximate to the leading edge 51 (line 5-5 is substantially at the same height as the floor 128 of the airfoil 28). The offset 37 in FIG. 5 is at the leading edge 51, but its position is only an example of possible positions and is not intended to limit the embodiments in any way. Offset 37 may be at any portion of the leading edge 51, preferably proximate to the leading edge 51, where the leading edge 51 transitions to the positive-pressure side wingtip cap portion 137.

[0026] Offset 37 and the step-down section 237 define a difference Z between the height Y and the height X. Offset 37 and the step-down section 237 can be formed during the manufacture of the airfoil 12 or can be formed after the airfoil 12 is formed. Further, the negative-pressure side wingtip cap portion 136 and the positive-pressure side wingtip cap portion 137 can be added or modified after the airfoil 12 has operated to enhance the flow from the wingtip cap 28 and the cavity 29.

[0027] According to some of the embodiments described above, it will be understood that this embodiment provides a way in which the wingtip cap configuration enables the flow of cooling fluid from a wingtip cap provided with ski-like wingtips. In addition, the wingtip cap enabled by this embodiment can allow for a tighter clearance between the blade and the surrounding fixed structure. As will be recognized by those skilled in the art, a tighter clearance causes more power gas to be moved against the airfoil, and in response, the rotor of the airfoil is moved to generate work.

[0028] As used throughout this specification and the claims, language that approximates is applicable to modify any quantitative expression that can vary within a reasonable degree without causing a change in the relevant basic function. Thus, values modified by terms such as "about," "approximately," and "substantially" are not limited to the exact values specified. In at least some instances, the language that approximates can correspond to the accuracy of the equipment for measuring the value. Here, as well as throughout this specification and the claims, limitations of ranges are combinable and / or replaceable, and such ranges are identified and include all sub-ranges subsumed therein, unless the context and language specifically dictate otherwise. "About" applied to a particular value of a range is applied to both end values and can indicate + / - 10% of the recited value, unless particularly dependent on the accuracy of the equipment for measuring the value.

[0029] All corresponding structures, materials, acts, and equivalents of the means-plus-function or step-plus-function elements in the following claims are intended to include any structure, material, or act for performing the function in combination with other claimed elements specifically recited for performing that function. The description of the present disclosure has been presented for purposes of illustration and description and is not intended to be exhaustive or to limit the disclosure to the forms disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the present disclosure. The embodiments have been chosen and described in order to best explain the principles of the disclosure and the practical application, and to enable others of ordinary skill in the art to understand the disclosure in various embodiments with various modifications as are suited to the particular use contemplated.

Description of Reference Numerals

[0030] 12 airfoil 20 concave positive pressure sidewall 22 convex negative pressure sidewall 28 outer wing tip, airfoil, wing tip cap 29 cavity 30 Internal cooling fluid flow path 34 Cooling fluid hole 36 Wing tip cap 37 Offset 50 Trailing edge 51 Leading edge 52 Opening 90 Turbo machine 100 Gas turbine (GT) system 102 Compressor 104 Combustor 105 Combustion region 106 Fuel nozzle assembly 108 Turbine 110 Rotor, turbine common shaft, rotating shaft 112 Stationary blade 114 Rotating blade 115 Stationary blade section 116 Rotor wheel 120 Radial outer end wall 122 Radial inner end wall 124 Casing 128 Floor 136 Negative pressure side wing tip cap part 137 Positive pressure side wing tip cap part 237 Step-down section F arrow L0 stage L1 stage L2 stage L3 stage X height Y height Z height change

Claims

[Claim 1] A blade (112), comprising: An airfoil (12) comprising: a floor (128), an outer tip (28) having a leading edge (51) and a trailing edge (50); and a concave pressure side wall (20) and a convex suction side wall (22) extending axially between corresponding leading and trailing edges (51) and radially between the floor (128) and the outer tip (28). Equipped with The airfoil portion (12) a tip cap (36) extending from the floor (128) of the outer tip (28) coextensive with the pressure side wall (20) and the suction side wall (22) and around each of the leading edge (51) and the trailing edge (50), the tip cap (36) comprising a squealer tip arrangement for reducing tip leakage and downstream mixing losses, the squealer tip arrangement comprising a suction side tip cap portion (136) and a pressure side tip cap portion (137). Furthermore, the suction side tip cap portion (136) and the pressure side tip cap portion (137) extend different distances above the floor (128), the tip cap (36) comprises an offset (37) providing a decreasing height change from a first height (Y) of the suction side tip cap portion (136) to a second height (X) of the pressure side tip cap portion (137), the offset (37) comprising a step-down section (237) from the first height (Y) of the suction side tip cap portion (136) to the second height (X) of the pressure side tip cap portion (137), and a flow of cooling fluid over the pressure side tip cap portion (137) of the tip cap (36) due to a pressure gradient; 1. A blade comprising: a blade having a step-down section of a pressure side tip cap portion extending substantially the entirety of the pressure side tip cap portion from the offset to an opening adjacent the trailing edge that allows the cooling fluid to exit the tip cap; and wherein the offset is selected from at least one of a sloped decrease in height change, a convexly curved decrease in height change, a concavely curved decrease in height change, or a combination thereof. **Claim 2**: The blade (112) according to claim 1, wherein the offset (37) is disposed close to the leading edge (51) such that the wing tip cap (36) transitions from the negative pressure side wing tip cap portion (136) to the step-down section (237) of the positive pressure side wing tip cap portion (137). **Claim 3** The blade (112) according to claim 1, wherein the height change is less than 100 mils (2.5 mm). **Claim 4** The blade (112) according to claim 1, wherein the height change is less than 50 mils (1.3 mm). **Claim 5** The blade (112) according to claim 1, wherein the height change is in the range from 10 mils (0.3 mm) to 50 mils (1.3 mm). **Claim 6** The blade (112) according to claim 1, wherein the offset (37) is formed in one of the positive pressure side wing tip cap portion (137), the leading edge (51), or the positive pressure side wing tip cap portion (137) at the leading edge (51), or the positive pressure side wing tip cap portion (137) adjacent to the leading edge (51). **Claim 7** The blade (112) according to claim 1, wherein the wing tip cap (36) has a cavity (29) between the floor (128), the negative pressure side wing tip cap portion (136), and the positive pressure side wing tip cap portion (137), and the blade (112) has at least one internal cooling fluid flow path (30) communicating with at least one corresponding cooling fluid hole in the floor (128). **Claim 8** A turbine (108) engine, wherein the turbine (108) engine comprises a blade (112) having an airfoil portion (12) according to any one of claims 1 to 7 The turbine (108) engine comprising the same.

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