Turbine Nozzle-To-Shroud Interface
A turbomachine includes a compressor, a combustor, and a turbine. The turbine includes a nozzle, a shroud, and an axial cavity defined between the nozzle and the shroud. The nozzle includes an inner platform, an outer platform spaced apart from the inner platform along a radial direction, and an airfoil extending therebetween. The outer platform extends axially between a forward sidewall and an aft sidewall. The nozzle also includes an aft hook radially outward of the outer platform and axially forward of the aft sidewall of the outer platform. The shroud includes a forward end adjacent to the aft sidewall of the outer platform. A sealing interface is positioned within the cavity, radially outward of the outer platform of the nozzle and axially forward of the aft sidewall of the outer platform of the nozzle.
The present disclosure generally relates to a turbine for a turbomachine. More particularly, this disclosure relates to a turbine nozzle-to-shroud interface for a turbomachine.
BACKGROUNDA gas turbine, such as an industrial, aircraft, or marine gas turbine generally includes, in serial flow order, a compressor, a combustor and a turbine. The turbine has multiple stages with each stage including a row of turbine nozzles and an adjacent row of turbine rotor blades disposed downstream from the turbine nozzles. The turbine nozzles are held stationary within the turbine and the turbine rotor blades rotate with a rotor shaft. The various turbine stages define a hot gas path through the turbine.
During operation, the compressor provides compressed air to the combustor. The compressed air is mixed with fuel and burned in a combustion chamber defined within the combustor to produce a high velocity stream of hot gas. The hot gas flows from the combustor into the hot gas path of the turbine via a turbine inlet. As the hot gas flows through each successive stage, kinetic energy from the high velocity hot gas is transferred to the rows of turbine rotor blades, thus causing the rotor shaft to rotate and produce mechanical work.
Turbine efficiency may be reduced when some of the hot gas flowing through the turbine hot gas path is lost before it can produce work. For example, the hot gas may be ingested into cavities between turbine components such as the turbine nozzles and adjacent shrouds.
BRIEF DESCRIPTIONAspects and advantages of the invention are set forth below in the following description, or may be obvious from the description, or may be learned through practice of the invention.
According to one embodiment, a gas turbine includes a central axis, the central axis of the gas turbine defines an axial direction, a radial direction perpendicular to the central axis, and a circumferential direction extending around the central axis. The gas turbine further includes a compressor, a combustor downstream of the compressor, and a turbine section downstream of the combustor. The turbine section includes a nozzle, a shroud, and a cavity defined by an axial clearance between the nozzle and the shroud. The nozzle includes an inner platform, an outer platform spaced apart from the inner platform along a radial direction, and an airfoil extending between the inner platform and the outer platform. The outer platform extends along an axial direction between a forward sidewall and an aft sidewall. The airfoil includes a leading edge at a forward end of the airfoil and a trailing edge at an aft end of the airfoil. The nozzle also includes an aft hook radially outward of the outer platform and axially forward of the aft sidewall of the outer platform. The shroud extends between a forward end and an aft end, the forward end of the shroud is adjacent to the aft sidewall of the outer platform. The shroud includes a forward hook configured to engage with the aft hook of the nozzle to support the nozzle. The turbine section further includes a sealing interface within the cavity. The sealing interface is radially outward of the outer platform of the nozzle and axially forward of the aft sidewall of the outer platform of the nozzle.
According to another embodiment, a turbine for a turbomachine is provided. The turbine includes a nozzle. The nozzle includes an inner platform and an outer platform spaced apart from the inner platform along a radial direction. The outer platform extends along an axial direction between a forward sidewall and an aft sidewall. The nozzle also includes an airfoil extending between the inner platform and the outer platform. The airfoil includes a leading edge at a forward end of the airfoil and a trailing edge at an aft end of the airfoil. The turbine also includes a shroud extending between a forward end and an aft end. The forward end of the shroud is adjacent to the aft sidewall of the outer platform. A cavity is defined by an axial clearance between the nozzle and the shroud. A sealing interface is positioned within the cavity. The sealing interface is radially outward of the outer platform of the nozzle and axially forward of the aft sidewall of the outer platform of the nozzle.
According to still another embodiment, a turbine for a turbomachine is provided. The turbine includes a nozzle. The nozzle includes an inner platform and an outer platform spaced apart from the inner platform along a radial direction. The outer platform extends along an axial direction between a forward sidewall and an aft sidewall. The nozzle also includes an airfoil extending between the inner platform and the outer platform. The airfoil includes a leading edge at a forward end of the airfoil and a trailing edge at an aft end of the airfoil. The airfoil also includes an aft hook radially outward of the outer platform and axially forward of the trailing edge of the airfoil. The turbine also includes a shroud extending between a forward end and an aft end. The forward end of the shroud is adjacent to the aft sidewall of the outer platform. The shroud includes a forward hook configured to engage with the aft hook of the nozzle to support the nozzle. A cavity is defined by an axial clearance between the nozzle and the shroud.
Those of ordinary skill in the art will better appreciate the features and aspects of such embodiments, and others, upon review of the specification.
A full and enabling disclosure of the present invention, including the best mode thereof to one skilled in the art, is set forth more particularly in the remainder of the specification, including reference to the accompanying figures, in which:
Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention.
As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. The term “radially” refers to the relative direction that is substantially perpendicular to an axial centerline of a particular component, the term “axially” refers to the relative direction that is substantially parallel and/or coaxially aligned to an axial centerline of a particular component, and the term “circumferentially” refers to the relative direction that extends around the axial centerline of a particular component.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and/or groups thereof.
Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present invention without departing from the scope or spirit thereof. For instance, features illustrated or described as part of one embodiment may be used on another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.
Although exemplary embodiments of the present invention will be described generally in the context of a turbine nozzle for a land based power generating gas turbine for purposes of illustration, one of ordinary skill in the art will readily appreciate that embodiments of the present invention may be applied to any style or type of gas turbine and are not limited to land based power generating gas turbines unless specifically recited in the claims.
Referring now to the drawings,
During operation, air 24 flows through the inlet section 12 and into the compressor 14 where the air 24 is progressively compressed, thus providing compressed air 26 to the combustor 16. At least a portion of the compressed air 26 is mixed with a fuel 28 within the combustor 16 and burned to produce combustion gases 30. The combustion gases 30 flow from the combustor 16 into the turbine 18, wherein energy (kinetic and/or thermal) is transferred from the combustion gases 30 to rotor blades, thus causing shaft 22 to rotate. The mechanical rotational energy may then be used for various purposes such as to power the compressor 14 and/or to generate electricity. The combustion gases 30 exiting the turbine 18 may then be exhausted from the gas turbine 10 via the exhaust section 20.
As shown in
The nozzles and shrouds of each stage 100, 200, 300 may include similar features as the nozzles and shrouds of any other stage 100, 200, 300. Accordingly, for the sake of clarity, the following description will refer primarily to the second stage nozzle 202 and second stage shroud 206 by way of example, it being understood that the first stage nozzle and shroud 102 and 106, and/or the third stage nozzle and shroud 302 and 306 may include any or all of the same features or similar features. For example, shroud 206 may extend between a forward end 226 and an aft end 228. In such embodiments, the forward end 226 of the shroud 206 may be adjacent to the aft sidewall 216 (
As illustrated in
As is generally understood in the art, the cavity 230 may define a flow path, e.g., along which hot gas may be ingested. It is also generally understood in the art that ingestion of hot gas into the cavity 230 may decrease the efficiency of the turbomachine 10, as less hot gas is available to perform work, e.g., in a downstream stage of the turbine 18. Accordingly, the positioning of the aft hook 222 of the nozzle 202 and other components defining the cavity 230 may advantageously provide a cavity 230 that defines a tortuous flow path. For example, as illustrated in
As illustrated in
Still with reference to
As is generally understood in the art, the airfoil 212 of the nozzle 202 may define a wake pressure zone W aft of the trailing edge 220. The wake pressure zone W may include circumferential pressure variations around the turbine section 18. For example, the axial extent of the wake pressure zone W of the exemplary nozzle 202 is illustrated in
In various embodiments of the present disclosure, the nozzle 202 and the shroud 206 may be constructed of any suitable materials. In some embodiments, the nozzle 202 may include a heat-resistant super alloy material, such as a nickel-based super alloy. In some embodiments, the shroud 206 may include a stainless steel material. Both the nickel-based super alloy and the stainless steel material are suitable for the nozzle 202 and the shroud 206, respectively. However, the nickel-based super alloy material of the nozzle 202 may have superior heat-resistant properties, yet the stainless steel material may be preferable for the shroud 206 since a substantial portion of the shroud 206 is not directly in the hot gas path. The shroud 206 may therefore be provided with various cooling features as are generally understood in the art. However, aspects of the present disclosure may advantageously shield the shroud 206 from the hot gas, e.g., the outer platform 210 may include a heat-resistant super alloy and may shield the shroud 206 from hot gas. Accordingly, the need for cooling features in the shroud 206 may be reduced.
As noted above, the foregoing example embodiments are described and illustrated with respect to a nozzle 202 and a shroud 206 of a second stage 200 of turbine 18 for simplicity and clarity only. The present disclosure is not limited to any particular stage 100, 200, and/or 300 of turbine 18, and various aspects of the example embodiments shown and described herein may be provided in any nozzle-to-shroud interface of a turbomachine.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other and examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.
Claims
1. A gas turbine, the gas turbine comprising a central axis, the central axis of the gas turbine defines an axial direction, a radial direction perpendicular to the central axis, and a circumferential direction extending around the central axis, the gas turbine further comprising:
- a compressor;
- a combustor downstream of the compressor;
- a turbine section downstream of the combustor, the turbine section comprising: a nozzle comprising: an inner platform; an outer platform spaced apart from the inner platform along a radial direction, the outer platform extending along an axial direction between a forward sidewall and an aft sidewall; an airfoil extending between the inner platform and the outer platform, the airfoil comprising a leading edge at a forward end of the airfoil and a trailing edge at an aft end of the airfoil; and an aft hook radially outward of the outer platform and axially forward of the aft sidewall of the outer platform; a shroud extending between a forward end and an aft end, the forward end of the shroud adjacent to the aft sidewall of the outer platform, the shroud comprising a forward hook, the forward hook of the shroud configured to engage with the aft hook of the nozzle to support the nozzle; a cavity defined by an axial clearance between the nozzle and the shroud; and a sealing interface within the cavity, the sealing interface radially outward of the outer platform of the nozzle and axially forward of the aft sidewall of the outer platform of the nozzle.
2. The gas turbine of claim 1, wherein the airfoil of the nozzle defines a wake pressure zone aft of the trailing edge, and the aft sidewall of the outer platform is positioned aft of the wake pressure zone.
3. The gas turbine of claim 1, wherein the cavity defines a tortuous flow path.
4. The gas turbine of claim 1, wherein the nozzle comprises a heat-resistant super alloy material.
5. The gas turbine of claim 1, wherein the shroud comprises a stainless steel material.
6. A turbine for a turbomachine, the turbine comprising:
- a nozzle comprising: an inner platform; an outer platform spaced apart from the inner platform along a radial direction, the outer platform extending along an axial direction between a forward sidewall and an aft sidewall; and an airfoil extending between the inner platform and the outer platform, the airfoil comprising a leading edge at a forward end of the airfoil and a trailing edge at an aft end of the airfoil;
- a shroud extending between a forward end and an aft end, the forward end of the shroud adjacent to the aft sidewall of the outer platform;
- a cavity defined by an axial clearance between the nozzle and the shroud; and
- a sealing interface within the cavity, the sealing interface radially outward of the outer platform of the nozzle and axially forward of the aft sidewall of the outer platform of the nozzle.
7. The turbine of claim 6, wherein the nozzle further comprises an aft hook positioned outward of the outer platform, the aft hook forward of the aft sidewall of the outer platform.
8. The turbine of claim 6, wherein the nozzle further comprises an aft hook positioned outward of the outer platform, the aft hook forward of the trailing edge of the airfoil.
9. The turbine of claim 6, wherein the airfoil of the nozzle defines a wake pressure zone aft of the trailing edge, and the aft sidewall of the outer platform is positioned aft of the wake pressure zone.
10. The turbine of claim 6, wherein the cavity defines a tortuous flow path.
11. The turbine of claim 6, wherein the nozzle comprises a heat-resistant super alloy material.
12. The turbine of claim 6, wherein the shroud comprises a stainless steel material.
13. A turbine for a turbomachine, the turbine comprising:
- a nozzle comprising: an inner platform; an outer platform spaced apart from the inner platform along a radial direction, the outer platform extending along an axial direction between a forward sidewall and an aft sidewall; an airfoil extending between the inner platform and the outer platform, the airfoil comprising a leading edge at a forward end of the airfoil and a trailing edge at an aft end of the airfoil; and an aft hook radially outward of the outer platform and axially forward of the trailing edge of the airfoil;
- a shroud extending between a forward end and an aft end, the forward end of the shroud adjacent to the aft sidewall of the outer platform, the shroud comprising a forward hook, the forward hook of the shroud configured to engage with the aft hook of the nozzle to support the nozzle; and
- a cavity defined by an axial clearance between the nozzle and the shroud.
14. The turbine of claim 13, further comprising a sealing interface between the nozzle and the shroud, the sealing interface radially outward of the outer platform of the nozzle and axially forward of the aft sidewall of the outer platform of the nozzle.
15. The turbine of claim 13, wherein the airfoil of the nozzle defines a wake pressure zone aft of the trailing edge, and the aft sidewall of the outer platform is positioned aft of the wake pressure zone.
16. The turbine of claim 13, wherein the cavity defines a tortuous flow path.
17. The turbine of claim 13, wherein the nozzle comprises a heat-resistant super alloy material.
18. The turbine of claim 13, wherein the shroud comprises a stainless steel material.
Type: Application
Filed: May 1, 2017
Publication Date: Nov 29, 2018
Inventors: Frederic Woodrow Roberts, JR. (Greer, SC), Sivaraman Vedhagiri (Greer, SC)
Application Number: 15/583,047