Rib in trailing edge cooling cavity of turbine airfoil

Abstract

A turbine component for a gas turbine engine includes an airfoil extending from an airfoil hub to an airfoil tip. The airfoil has a pressure side wall and a suction side wall that meet at a leading edge and a trailing edge. A trailing edge cooling cavity is defined in an airfoil interior between the pressure side wall and the suction side wall. The trailing edge cooling cavity includes a rib having a rib root connected to one of the pressure side wall and suction side wall, a rib tip positioned into the airfoil interior without reaching the other one of the pressure side wall and the suction side wall, and a rib body extending a non-zero distance from the rib root to the rib tip and extending continuously from the airfoil hub to the airfoil tip as a single solid piece.

Description

Docket No. 2024PF00291RIB IN TRAILING EDGE COOLING CAVITY OF TURBINE AIRFOIL BACKGROUND

[0001] A gas turbine engine typically includes a compressor section, a turbine section, and a combustion section disposed therebetween. The compressor section typically includes multiple stages of rotating compressor blades and stationary compressor vanes to produce compressed air. The combustion section typically includes a plurality of combustors to produce hot working fluid by combusting mixture of the compressed air and fuel. The turbine section typically includes multiple stages of rotating turbine blades and stationary turbine vanes to expand the hot working fluid and to convert fluid energy to mechanical energy.

[0002] The turbine components are often manufactured by a casting process. The casting process is prone to defects which are detrimental to the life of the turbine components. It is desirable to improve the casting process to mitigate the defects.BRIEF SUMMARY

[0003] In one aspect, a turbine component for a gas turbine engine is provided. The turbine component includes an airfoil that extends from an airfoil hub to an airfoil tip. The turbine component also includes an airfoil wall that defines an airfoil interior. The airfoil wall includes a pressure side wall and a suction side wall that meet at a leading edge and a trailing edge. The turbine component also includes a trailing edge cooling cavity defined in the airfoil interior of a trailing edge region. The turbine component also includes a rib disposed in the trailing edge cooling cavity. The rib includes a rib root that is connected to one of the pressure side wall and the suction side wall, a rib tipDocket No. 2024PF00291that is positioned into the airfoil interior without reaching the other one of the pressure side wall and the suction side wall, a rib body that extends a non-zero distance from the rib root to the rib tip and extends continuously from the airfoil hub to the airfoil tip, the rib body being a single solid piece.

[0004] In one aspect, a turbine component for a gas turbine engine is provided. The turbine component includes an airfoil that extends from an airfoil hub to an airfoil tip. The turbine component also includes an airfoil wall that defines an airfoil interior. The airfoil wall includes a pressure side wall and a suction side wall that meet at a leading edge and a trailing edge. The turbine component also includes a trailing edge cooling cavity defined in the airfoil interior of a trailing edge region. The turbine component also includes a pressure side rib that extends from the pressure side wall into the trailing edge cooling cavity. The turbine component also includes a suction side rib that extends from the suction side wall into the trailing edge cooling cavity. Each rib of the pressure side rib and the suction side rib includes a rib root that is connected to one of the pressure side wall and the suction side wall, a rib tip that is positioned into the airfoil interior without reaching the other one of the pressure side wall and suction side wall, and a rib body that extends a non-zero distance from the rib root to the rib tip and extends continuously from the airfoil hub to the airfoil tip, the rib body being a single solid piece.BRIEF DESCRIPTION OF THE DRAWINGS

[0005] To easily identify the discussion of any particular element or act, the most significant digit or digits in a reference number refer to the figure number in which that element is first introduced.

[0006] FIG. 1 is a longitudinal cross-sectional view of a gas turbine engine taken along a plane that contains a longitudinal axis or central axis.

[0007] FIG. 2 illustrates a perspective view of a turbine component for use with the gas turbine engine of FIG. 1.Docket No. 2024PF00291

[0008] FIG. 3 illustrates a cross-sectional view of an airfoil of the turbine component of FIG. 2.

[0009] FIG. 4 illustrates a perspective view of a trailing edge region of the airfoil of FIG. 3 with the suction side wall removed.

[0010] FIG. 5 illustrates a cross-sectional view of the trailing edge region of the airfoil of FIG. 3 near an airfoil hub along a section plane 5-5 of FIG. 2.

[0011] FIG. 6 illustrates a cross-sectional view of the trailing edge region of the airfoil of FIG. 3 at a taper transition region along a section plane 6-6 of FIG. 2.

[0012] FIG. 7 illustrates a cross-sectional view of the trailing edge region of the airfoil of FIG. 3 near an airfoil tip along a section plane 7-7 of FIG. 2.DETAILED DESCRIPTION

[0013] Before any embodiments of the invention are explained in detail, it is to be understood that the invention is not limited in its application to the details of construction and the arrangement of components set forth in this description or illustrated in the following drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting.

[0014] Various technologies that pertain to systems and methods will now be described with reference to the drawings, where like reference numerals represent like elements throughout. The drawings discussed below, and the various embodiments used to describe the principles of the present disclosure in this patent document are by way of illustration only and should not be construed in any way to limit the scope of the disclosure. Those skilled in the art will understand that the principles of the present disclosure may be implemented in any suitably arranged apparatus. It is to beDocket No. 2024PF00291understood that functionality that is described as being carried out by certain system elements may be performed by multiple elements. Similarly, for instance, an element may be configured to perform functionality that is described as being carried out by multiple elements. The numerous innovative teachings of the present application will be described with reference to exemplary non-limiting embodiments.

[0015] It should be understood that the words or phrases used herein should be construed broadly, unless expressly limited in some examples. For example, the terms “including”, “having”, and “comprising”, as well as derivatives thereof, mean inclusion without limitation. The singular forms “a”, “an”, and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. Further, the term “and / or” as used herein refers to and encompasses any and all possible combinations of one or more of the associated listed items. The term “or” is inclusive, meaning and / or, unless the context clearly indicates otherwise. The phrases “associated with” and “associated therewith” as well as derivatives thereof, may mean to include, be included within, interconnect with, contain, be contained within, connect to or with, couple to or with, be communicable with, cooperate with, interleave, juxtapose, be proximate to, be bound to or with, have, have a property of, or the like. Furthermore, while multiple embodiments or constructions may be described herein, any features, methods, steps, components, etc. described with regard to one embodiment are equally applicable to other embodiments absent a specific statement to the contrary.

[0016] Although the terms “first”, “second”, “third” and so forth may be used herein to refer to various elements, information, functions, or acts, these elements, information, functions, or acts should not be limited by these terms. Rather these numeral adjectives are used to distinguish different elements, information, functions or acts from each other. For example, a first element, information, function, or act could be termed a second element, information, function, or act, and, similarly, a second element, information, function, or act could be termed a first element, information, function, or act, without departing from the scope of the present disclosure.Docket No. 2024PF00291

[0017] In the description, the terms “axial” or “axially” refer to a direction along a longitudinal axis of a gas turbine engine. The terms “radial” or “radially” refer to a direction perpendicular to the longitudinal axis of the gas turbine engine. The terms “downstream” or “aft” refer to a direction along a flow direction. The terms “upstream” or “forward” refer to a direction against the flow direction.

[0018] In addition, the term “adjacent to" may mean that an element is relatively near to but not in contact with a further element or that the element is in contact with the further portion, unless the context clearly indicates otherwise. Further, the phrase “based on” is intended to mean “based, at least in part, on” unless explicitly stated otherwise. Terms “about” or “substantially” or like terms are intended to cover variations in a value that are within normal industry manufacturing tolerances for that dimension. If no industry standard is available, a variation of twenty percent would fall within the meaning of these terms unless otherwise stated.

[0019] FIG. 1 illustrates an example of a gas turbine engine 100 including a compressor section 102, a combustion section 104, and a turbine section 106 arranged along a central axis 108. The compressor section 102 includes a plurality of compressor stages 110 with each compressor stage 110 including a set of stationary compressor vanes 112 or adjustable guide vanes and a set of rotating compressor blades 114. A rotor 116 supports the rotating compressor blades 114 for rotation about the central axis 108 during operation. In some constructions, a single one-piece rotor 116 extends the length of the gas turbine engine 100 and is supported for rotation by a bearing at either end. In other constructions, the rotor 116 is assembled from several separate spools that are attached to one another or may include multiple disk sections that are attached via a bolt or plurality of bolts.

[0020] The compressor section 102 is in fluid communication with an inlet section 118 to allow the gas turbine engine 100 to draw atmospheric air into the compressor section 102. During operation of the gas turbine engine 100, the compressor section 102 draws in atmospheric air and compresses that air for delivery to the combustion section 104.Docket No. 2024PF00291The illustrated compressor section 102 is an example of one compressor section 102 with other arrangements and designs being possible.

[0021] In the illustrated construction, the combustion section 104 includes a plurality of separate combustors 120 that each operate to mix a flow of fuel with the compressed air from the compressor section 102 and to combust that air-fuel mixture to produce a flow of high temperature, high pressure combustion gas or exhaust gas 122. Of course, many other arrangements of the combustion section 104 are possible.

[0022] The turbine section 106 includes a plurality of turbine stages 124 with each turbine stage 124 including a number of stationary turbine vanes 126 and a number of rotating turbine blades 128. The turbine stages 124 are arranged to receive the exhaust gas 122 from the combustion section 104 at a turbine inlet 130 and expand that gas to convert thermal and pressure energy into rotating or mechanical work. The turbine section 106 is connected to the compressor section 102 to drive the compressor section 102. For gas turbine engines 100 used for power generation or as prime movers, the turbine section 106 is also connected to a generator, pump, or other device to be driven. As with the compressor section 102, other designs and arrangements of the turbine section 106 are possible.

[0023] An exhaust portion 132 is positioned downstream of the turbine section 106 and is arranged to receive the expanded flow of exhaust gas 122 from the final turbine stage 124 in the turbine section 106. The exhaust portion 132 is arranged to efficiently direct the exhaust gas 122 away from the turbine section 106 to assure efficient operation of the turbine section 106. Many variations and design differences are possible in the exhaust portion 132. As such, the illustrated exhaust portion 132 is but one example of those variations.

[0024] A control system 134 is coupled to the gas turbine engine 100 and operates to monitor various operating parameters and to control various operations of the gas turbine engine 100. In preferred constructions the control system 134 is typically microprocessor based and includes memory devices and data storage devices for collecting, analyzing, and storing data. In addition, the control system 134 provides output data toDocket No. 2024PF00291various devices including monitors, printers, indicators, and the like that allow users to interface with the control system 134 to provide inputs or adjustments. In the example of a power generation system, a user may input a power output set point and the control system 134 may adjust the various control inputs to achieve that power output in an efficient manner.

[0025] The control system 134 can control various operating parameters including, but not limited to variable inlet guide vane positions, fuel flow rates and pressures, engine speed, valve positions, generator load, and generator excitation. Of course, other applications may have fewer or more controllable devices. The control system 134 also monitors various parameters to assure that the gas turbine engine 100 is operating properly. Some parameters that are monitored may include inlet air temperature, compressor outlet temperature and pressure, combustor outlet temperature, fuel flow rate, generator power output, bearing temperature, and the like. Many of these measurements are displayed for the user and are logged for later review should such a review be necessary.

[0026] FIG. 2 illustrates a perspective view of a turbine component 200. The turbine component 200 or similar turbine components may be used in the gas turbine engine 100. The turbine component 200 may be the stationary turbine vane 126 or the rotating turbine blade 128. For illustration purpose, the turbine component 200 shown in FIG. 2 is the stationary turbine vane 126. It is understood that the configurations illustrated in FIG. 2 can be incorporated to the rotating turbine blade 128.

[0027] The turbine component 200 has an airfoil 202 that extends from an airfoil hub 204 to an airfoil tip 206 to define a span in a radial direction R. The span is a distance between the airfoil hub 204 and the airfoil tip 206. The airfoil hub 204 is connected to an outer platform 208. The airfoil tip 206 is connected to an inner platform 210.

[0028] FIG. 3 illustrates a cross-sectional view of the airfoil 202. The airfoil wall 212 defines an airfoil interior 302. The airfoil wall 212 includes a pressure side wall 304 and a suction side wall 306 that meet at a leading edge 308 and a trailing edge 310. The airfoil interior 302 is defined between the pressure side wall 304 and the suction sideDocket No. 2024PF00291wall 306. A plurality of coolant exit slots 312 are disposed at the trailing edge 310 along the radial direction. The minimum thickness of the airfoil wall 212 is between 0.5 mm to 1.5 mm near the trailing edge 310.

[0029] The airfoil 202 has a front rib 314 and a rear rib 316 that connect the pressure side wall 304 and the suction side wall 306 and extend along the radial direction. The front rib 314 and the rear rib 316 divide the airfoil interior 302 into a leading edge region 318, a middle region 320, and a trailing edge region 322 along an axial direction X parallel to the central axis 108 of the gas turbine engine 100. The leading edge region 318 is adjacent to the leading edge 308 and defined between the leading edge 308 and the front rib 314. The middle region 320 is defined between the front rib 314 and the rear rib 316. The trailing edge region 322 is adjacent to the trailing edge 310 and defined between the rear rib 316 and the trailing edge 310. Cooling cavities are formed in each of the leading edge region 318, the middle region 320, and the trailing edge region 322 to cool the turbine component 200.

[0030] A trailing edge cooling cavity 324 is defined in the airfoil interior 302 of the trailing edge region 322 and between the pressure side wall 304 and the suction side wall 306. A plurality of ribs 326 are disposed in the trailing edge cooling cavity 324 and spaced apart from each other in the axial direction toward the trailing edge 310. The plurality of ribs 326 extend from one of the pressure side wall 304 and the suction side wall 306 into the trailing edge cooling cavity 324 without reaching the other one of the pressure side wall 304 and the suction side wall 306. Each rib 326 of the plurality of ribs 326 has a rib root 328, a rib tip 330, and a rib body 332 that extends a non-zero distance from the rib root 328 to the rib tip 330. The rib root 328 is connected to one of the pressure side wall 304 and the suction side wall 306. The rib tip 330 is positioned into the trailing edge cooling cavity 324 without reaching the other one of the pressure side wall 304 and the suction side wall 306. A height of the rib 326 is defined as a distance between a mid-point of the rib root 328 and a mid-point of the rib tip 330. The height of the rib 326 decreases along the axial direction toward the trailing edge 310. In other constructions, the trailing edge cooling cavity 324 may have a single rib 326 that extends from one of the pressure side wall 304 and the suction side wall 306 into theDocket No. 2024PF00291trailing edge cooling cavity 324 without reaching the other one of the pressure side wall 304 and the suction side wall 306. In other constructions, the height of the rib 326 may be constant along the axial direction toward the trailing edge 310.

[0031] The plurality of ribs 326 include a plurality of pressure side ribs 334 and a plurality of suction side ribs 336. The plurality of pressure side ribs 334 extend from the pressure side wall 304 into the trailing edge cooling cavity 324 without reaching the suction side wall 306. The plurality of suction side ribs 336 extend from the suction side wall 306 into the trailing edge cooling cavity 324 without reaching the pressure side wall 304. The pressure side ribs 334 and the suction side ribs 336 are arranged alternatively from the pressure side wall 304 and the suction side wall 306 to form a zigzag cooling path 338 toward the trailing edge 310 for coolant 340. The zigzag cooling path 338 is in fluid communication with the coolant exit slots 312 to exit the coolant 340 from the turbine component 200. In other constructions, the trailing edge cooling cavity 324 may have a single pressure side rib 334 and a single suction side rib 336 that extend from one of the pressure side wall 304 and the suction side wall 306 into the trailing edge cooling cavity 324 without reaching the other one of the pressure side wall 304 and the suction side wall 306.

[0032] Each rib body 332 has a first planar side surface and a second planar side surface between the rib root 328 and the rib tip 330. Each rib body 332 has a tapered shape from the rib root 328 to the rib tip 330. The taper is defined as a change of a width of the rib root 328 and the rib tip 330 divided by the height of the rib 326. The width of the rib root 328 is larger than a width of the rib tip 330. The width of the rib root 328 is a distance between one side of the rib root 328 to the other side of the rib 326 in the axial direction. The width of the rib tip 330 is a distance between one side the rib tip 330 to the other side of the rib tip 330 in the axial direction.

[0033] Each rib 326 extends from one of the pressure side wall 304 and the suction side wall 306 at an oblique angle with respect to one of the pressure side wall 304 and the suction side wall 306. The oblique angle is between 20 degrees to 89 degrees. The oblique angle has the same direction from the airfoil hub 204 to the airfoil tip 206. EachDocket No. 2024PF00291rib 326 has a constant oblique angle from the airfoil hub 204 to the airfoil tip 206. The plurality of ribs 326 have the same oblique angle with each other. In other constructions, the plurality of ribs 326 or some of the plurality of ribs 326 may extend perpendicularly to the one of the pressure side wall 304 and the suction side wall 306. In other constructions, the oblique angle of the plurality of ribs 326 or some of the plurality of ribs 326 may be less than 20 degrees. In other construction, the oblique angle of the plurality of ribs 326 or some of the plurality of ribs 326 may change between the airfoil hub 204 and the airfoil tip 206. In other construction, the oblique angle of the plurality of ribs 326 or some of the plurality of ribs 326 may be different from each other.

[0034] FIG. 4 illustrates a perspective view of the trailing edge region 322. For illustration purpose, the suction side wall 306 is removed in FIG. 4.

[0035] The plurality of pressure side ribs 334 extend continuously from the airfoil hub 204 to the airfoil tip 206 without interspaces therebetween. Each pressure side rib 334 of the pressure side ribs 334 is a single solid piece from the airfoil hub 204 to the airfoil tip 206.

[0036] The cross-sectional area of each pressure side rib 334 decreases from the airfoil hub 204 to a taper transition region and increases from the taper transition region to the airfoil tip 206. The width of the rib root 328 of each pressure side rib 334 decreases from the airfoil hub 204 to the taper transition region and increases from the taper transition region to the airfoil tip 206. The width of the rib tip 330 of each pressure side rib 334 decreases from the airfoil hub 204 to the taper transition region and increases from the taper transition region to the airfoil tip 206. The taper transition region is defined as a span region between 30% span to 70% span from the airfoil hub 204 to the airfoil tip 206. In other constructions, each pressure side rib 334 or some of the pressure side ribs 334 may have a constant cross-sectional area along the radial direction between the airfoil hub 204 and the airfoil tip 206. In other construction, the taper transition region may be different than the span region of 30% span to 70% span from the airfoil hub 204 to the airfoil tip 206, such as 20% span to 80% span from the airfoil hub 204 toDocket No. 2024PF00291the airfoil tip 206, or 40% span to 60% span from the airfoil hub 204 to the airfoil tip 206.

[0037] The suction side ribs 336 have the similar configuration as the pressure side ribs 334 and are not described in detail herewith.

[0038] FIG. 5 illustrates a cross-sectional view of the trailing edge region 322 near the airfoil hub 204 along a section 5-5 of FIG. 2. FIG. 6 illustrates a cross-sectional view of the trailing edge region 322 at the taper transition region along a section 6-6 of FIG. 2. FIG. 7 illustrates a cross-sectional view of the trailing edge region 322 near the airfoil tip 206 along a section 7-7 of FIG. 2.

[0039] With references to FIG. 5, FIG. 6, and FIG. 7, the width of the rib root 328 and the width of the rib tip 330 at the taper transition region are smaller than the width of the rib root 328 and the width of the rib tip 330 near the airfoil hub 204 and near the airfoil tip 206 to define a larger cooling passage of the zigzag cooling path 338 at the taper transition region. The width of the rib root 328 and the width of the rib tip 330 near the airfoil hub 204 are larger than the width of the rib root 328 and the width of the rib tip 330 near the airfoil tip 206.

[0040] The hot gas path turbine components 200 are typically cast out of nickel-based superalloys using a casting core. The casting process is prone to defects which are detrimental to the life of the turbine component 200. The heavy taper of the ribs 326 mitigates the defects which improves the structure strength of the turbine component 200.

[0041] The casting core represents the hollow cooling flow passages in the airfoil interior 302. The continuous solid piece ribs 326 from the airfoil hub 204 to the front rib 314 represent a continuous solid piece casting core from the airfoil hub 204 to the airfoil tip 206 which mitigates casting core breakage and inclusions. The deceasing cross-sectional area of the ribs 326 from the airfoil hub 204 and the airfoil tip 206 to the taper transition region represents an increasing cross-sectional area of the casting core from the airfoil hub 204 and the airfoil tip 206 to the taper transition region which aidsDocket No. 2024PF00291in core injection and helps in core yield as well as strengthening the casting core. The minimum thickness of 0.5 mm to 1.5 mm of the airfoil wall 212 near the trailing edge 310 represents a minimum thickness of 1.5 mm to 2.0 mm of the casting core which reduces the risk of core brakeage. The oblique angle of the ribs 326 improves the manufacture process and helps in solidification feeding during casting. The large mass of the ribs 326 provides enough heat to mitigate chill grain during casting. The tapered shapes of the ribs 326 reduces shrinkage porosity and improves the strength of the ribs 326.

[0042] In operation of the gas turbine engine 100, the coolant 340 enters the trailing edge cooling cavity 324 and flows through the zigzag cooling path 338 to cool the turbine component 200 and exits the turbine component 200 from the coolant exit slots 312. The continuous solid ribs 326 from the airfoil hub 204 to the airfoil tip 206 without interspaces therebetween provide a large heat transfer surface area and thus aid in improving internal convective heat transfer mechanism. The continuous solid ribs 326 from the airfoil hub 204 to the airfoil tip 206 without interspaces therebetween also provide a large cross-sectional area and thus aid in improving internal conductive heat transfer mechanism. The combination of an increased internal convective and conductive heat transfer results in a reduction of internal heat transfer resistance which leads to an improvement in cooling effect of the coolant 340.

[0043] Although an exemplary embodiment of the present disclosure has been described in detail, those skilled in the art will understand that various changes, substitutions, variations, and improvements disclosed herein may be made without departing from the spirit and scope of the disclosure in its broadest form.

[0044] None of the description in the present application should be read as implying that any particular element, step, act, or function is an essential element, which must be included in the claim scope: the scope of patented subject matter is defined only by the allowed claims. Moreover, none of these claims are intended to invoke a means plus function claim construction unless the exact words "means for" are followed by a participle.

Claims

1. Docket No. 2024PF00291CLAIMSWhat is claimed is:

1. A turbine component for a gas turbine engine, the turbine component comprising:an airfoil that extends from an airfoil hub to an airfoil tip;an airfoil wall that defines an airfoil interior, the airfoil wall comprising a pressure side wall and a suction side wall that meet at a leading edge and a trailing edge;a trailing edge cooling cavity defined in the airfoil interior of a trailing edge region; anda rib disposed in the trailing edge cooling cavity, the rib comprising:a rib root that is connected to one of the pressure side wall and the suction side wall;a rib tip that is positioned into the airfoil interior without reaching the other one of the pressure side wall and the suction side wall; anda rib body that extends a non-zero distance from the rib root to the rib tip and extends continuously from the airfoil hub to the airfoil tip, the rib body being a single solid piece.

2. The turbine component of claim 1, wherein the rib body has a tapered shape from the rib root to the rib tip.

3. The turbine component of claim 1, wherein the rib extends from one of the pressure side wall and the suction side wall at an oblique angle with respect to the one of the pressure side wall and the suction side wall.

4. The turbine component of claim 3, wherein the oblique angle is between 20 degrees to 89 degrees.

5. The turbine component of claim 3, wherein the oblique angle is constant from the airfoil hub to the airfoil tip.Docket No. 2024PF002916. The turbine component of claim 1, wherein a width of the rib root decreases from the airfoil hub to a taper transition region and increases from the taper transition region to the airfoil tip.

7. The turbine component of claim 1, wherein a width of the rib tip decreases from the airfoil hub to a taper transition region and increases from the taper transition region to the airfoil tip.

8. The turbine component of claim 1, wherein a width of the rib root of each rib at the airfoil hub is larger than a width of the rib root of the each rib at the airfoil tip.

9. The turbine component of claim 1, wherein a width of the rib tip of each rib at the airfoil hub is larger than a width of the rib tip of the each rib at the airfoil tip.

10. The turbine component of claim 1, further comprising a plurality of ribs that are disposed in the trailing edge region and spaced apart from each other toward the trailing edge, wherein the plurality of ribs extend alternatively from the pressure side wall and the suction side wall.

11. A turbine component for a gas turbine engine, the turbine component comprising:an airfoil that extends from an airfoil hub to an airfoil tip;an airfoil wall that defines an airfoil interior, the airfoil wall comprising a pressure side wall and a suction side wall that meet at a leading edge and a trailing edge;a trailing edge cooling cavity defined in the airfoil interior of a trailing edge region;a pressure side rib that extends from the pressure side wall into the trailing edge cooling cavity; anda suction side rib that extends from the suction side wall into the trailing edge cooling cavity,wherein each rib of the pressure side rib and the suction side rib comprises:a rib root that is connected to one of the pressure side wall and the suction side wall;Docket No. 2024PF00291a rib tip that is positioned into the airfoil interior without reaching the other one of the pressure side wall and suction side wall; anda rib body that extends a non-zero distance from the rib root to the rib tip and extends continuously from the airfoil hub to the airfoil tip, the rib body being a single solid piece.

12. The turbine component of claim 11, wherein the rib body has a tapered shape from the rib root to the rib tip.

13. The turbine component of claim 11 , wherein each rib extends from one of the pressure side wall and the suction side wall at an oblique angle with respect to the one of the pressure side wall and the suction side wall.

14. The turbine component of claim 13, wherein the oblique angle is between 20 degrees to 89 degrees.

15. The turbine component of claim 13, wherein the oblique angle is constant from the airfoil hub to the airfoil tip.

16. The turbine component of claim 11, wherein a width of the rib root of each rib decreases from the airfoil hub to a taper transition region and increases from the taper transition region to the airfoil tip.

17. The turbine component of claim 11, wherein a width of the rib tip of each rib decreases from the airfoil hub to a taper transition region and increases from the taper transition region to the airfoil tip.

18. The turbine component of claim 11, wherein a width of the rib root of each rib at the airfoil hub is larger than a width of the rib root of the each rib at the airfoil tip.

19. The turbine component of claim 11, wherein a width of the rib tip of each rib at the airfoil hub is larger than a width of the rib tip of the each rib at the airfoil tip.

20. The turbine component of claim 11, further comprising a plurality of pressure side ribs and a plurality of suction side ribs that are disposed in the trailing edge regionDocket No. 2024PF00291and spaced apart from each other toward the trailing edge, wherein the plurality of pressure side ribs and the plurality of suction side ribs extend alternatively from the pressure side wall and the suction side wall.

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