Patent Application
20200141276
The present disclosure relates to a gas turbine engine and, more specifically, to segmented turbine shrouds and to systems for reducing leakage between adjacent shroud segments using lapped seal joints.
A typical gas turbine engine, as may be used to propel an aircraft, generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet, or intake, is located at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a fan, a compressor, a combustion chamber, and a turbine. It will be readily apparent to those skilled in the art that additional components may also be included in the engine, such as low-pressure and high-pressure compressors and low-pressure and high-pressure turbines. This, however, is not an exhaustive list.
The compressor and turbine generally include rows of airfoils that are stacked axially in stages. Each stage includes a row of circumferentially spaced stator vanes that are affixed to an outer casing and a row of rotor blades that rotate about a center shaft or axis of the turbine engine. A multi-stage, low-pressure turbine follows the multi-stage high-pressure turbine and is typically joined by a second shaft to a fan disposed upstream from the compressor in a typical turbo-fan aircraft engine configuration used for powering an aircraft in flight.
A typical gas turbine engine utilizes a high-pressure turbine and a low-pressure turbine to maximize extraction of energy from high temperature combustion gas.
The turbine section typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The blades are circumferentially distributed on a rotor, causing rotation of the internal shaft. The internal shaft is connected to the rotor and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades. This powers the compressor during operation and subsequently drives the turbine. As the combustion gas flows downstream through the turbine stages, energy is extracted from the combustion gas, and the pressure of the combustion gas is reduced.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases that flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high-pressure turbine first receives the hot combustion gases from the combustor and includes a row of stator nozzles directing the combustion gases downstream through a row of high-pressure turbine rotor blades that extend radially outward from a supporting rotor disk. The stator nozzles turn the hot combustion gas in a manner to maximize extraction at the adjacent downstream turbine blades. In a two-stage turbine, a second row of stator nozzles is positioned downstream of the first stage blades, followed in turn by a row of second stage rotor blades that extend radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy.
During the operation of the gas turbine engine, it is desirable to minimize thermally induced deformation of the outer casing through the turbine section of the engine. In some cases, turbine shrouds are used to isolate the outer casing from heat produced by the hot combustion gases flowing through the turbine. The turbine shrouds are connected to the engine casing, via hangers, to provide an outer flow boundary for the combustion gas, which limits the high-temperature combustion gas from adversely affecting the casing. The shroud extends circumferentially to form a ring shape and may be formed of a plurality of circumferentially extending shroud segments.
The rotation of the turbine blade causes the combustion gas to move radially outward toward the turbine shrouds. If the combustion gas leaks through the joints between adjacent shroud segments, energy losses may occur, leading to sub-optimal performance of the gas turbine. To combat these leakages, multiple small seals are positioned between shroud segments (one of which is shown in
The circumferential end face 14 of the shroud segment 2 includes seal slots 16 within which individual seals are installed. The circumferential end face 41 of the hanger 40 includes additional seal slots 49 within which individual seals are installed. The seals (not shown) are typically thin metal strips that are inserted individually between each shroud segment as the shroud segments are installed.
As may be imagined, it is often difficult to maintain the seals within the slots of a first shroud segment, while positioning the adjacent shroud segment. If one of the seals falls out, a technician must search for the displaced seal and re-install it. This is a time-consuming and tedious process, especially since some of the seals may be very short (around 1 inch). Moreover, because the seals are made of thin metal, they are prone to burn through at higher operating temperatures, which can negatively impact the durability of the turbine shroud as well as the performance of the gas turbine.
It would be desirable to overcome these and other deficiencies with improved turbine shroud segments for gas turbine engines. More specifically, it would be desirable to eliminate some of or all the seals between circumferentially adjacent shroud segments, while maintaining a sealed joint between these shroud segments.
According to a first aspect of the present disclosure, a turbine shroud is disposed about a plurality of rotating turbine blades. The turbine shroud includes a plurality of arcuate shroud segments sealingly engaged with one another. Each arcuate shroud segment includes a bottom panel proximate to the plurality of blades. The bottom panel has a radially inner surface and a radially outer surface, a first end, and a second end. The first end includes a first end face, a first end cutback face circumferentially spaced from the first end face, and an intermediate surface extending between the first end face and the end cutback face to define a recess. The second end includes a second end face, a second end cutback face circumferentially spaced from the second end face, and a projecting surface extending between the second end face and the second end cutback face. The projecting surface of each shroud segment overlaps the recess of a circumferentially adjacent shroud segment to produce a sealed joint.
According to another aspect of the present disclosure, a turbine engine includes a rotor comprising a plurality of stationary nozzles arranged in a first stage and a second stage; and a shroud disposed between the first stage and the second stage of the plurality of stationary nozzles. The turbine shroud includes a plurality of arcuate shroud segments sealingly engaged with one another. Each arcuate shroud segment includes a bottom panel proximate to the plurality of blades. The bottom panel has a radially inner surface and a radially outer surface, a first end, and a second end. The first end includes a first end face, a first end cutback face circumferentially spaced from the first end face, and an intermediate surface extending between the first end face and the end cutback face to define a recess. The second end includes a second end face, a second end cutback face circumferentially spaced from the second end face, and a projecting surface extending between the second end face and the second end cutback face. The projecting surface of each shroud segment overlaps the recess of a circumferentially adjacent shroud segment to produce a sealed joint.
The specification, directed to one of ordinary skill in the art, sets forth a full and enabling disclosure of the present system and method, including the best mode of using the same. The specification refers to the appended figures, in which:
and
Reference will now be made in detail to various embodiments of the present disclosure, 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 disclosure.
To clearly describe the current turbine shroud segments, certain terminology will be used to refer to and describe relevant machine components within the scope of this disclosure. To the extent possible, common industry terminology will be used and employed in a manner consistent with the accepted meaning of the terms. Unless otherwise stated, such terminology should be given a broad interpretation consistent with the context of the present application and the scope of the appended claims. Those of ordinary skill in the art will appreciate that often a particular component may be referred to using several different or overlapping terms. What may be described herein as being a single part may include and be referenced in another context as consisting of multiple components. Alternatively, what may be described herein as including multiple components may be referred to elsewhere as a single integrated part.
In addition, several descriptive terms may be used regularly herein, as described below. 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.
As used herein, “downstream” and “upstream” are terms that indicate a direction relative to the flow of a fluid, such as the working fluid through the turbine engine. The term “downstream” corresponds to the direction of flow of the fluid, and the term “upstream” refers to the direction opposite to the flow (i.e., the direction from which the fluid flows). The terms “forward” and “aft,” without any further specificity, refer to relative position, with “forward” being used to describe components or surfaces located toward the front (or compressor) end of the engine or toward the inlet end of the combustor, and “aft” being used to describe components located toward the rearward (or turbine) end of the engine or toward the outlet end of the combustor. The term “inner” is used to describe components in proximity to the turbine shaft, while the term “outer” is used to describe components distal to the turbine shaft.
It is often required to describe parts that are at differing radial, axial and/or circumferential positions. As shown in
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting. 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, not limitation. In fact, it will be apparent to those skilled in the art that modifications and variations can be made 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 disclosure covers such modifications and variations as come within the scope of the appended claims and their equivalents.
Although exemplary embodiments of the present disclosure will be described generally in the context of turbine shrouds for aviation-powering gas turbines for purposes of illustration, one of ordinary skill in the art will readily appreciate that embodiments of the present disclosure may be applied to land-based power-generating gas turbines and are not limited to turbine components for gas turbines in a particular field, unless specifically recited in the claims.
Referring now to the present turbine shroud segments,
The hanger 140 mounts to the casing (not shown) by means of tabs 147, 148 and to adjacent stationary nozzles 60 by means of oppositely disposed ribs 143 that are surrounded by hooks 62 projecting from the surface of the stationary nozzles 60. The stationary nozzles 60 are representative of a plurality of circumferentially spaced stationary nozzles that are grouped in stages. It should be understood that the shroud made from the present shroud segments 100 (or 200) may be positioned between any pair of adjacent stages of stationary nozzles 60.
The turbine shroud segment 100 includes a bottom panel 110, a first leg 120 extending radially outward from the bottom panel 110 along the circumferential length of the shroud segment 100, and a second leg 130 extending radially outward from the bottom panel 110 along the circumferential length of the shroud segment 100 opposite the first leg 120. The intersection of the first leg 120 and the bottom panel 110 defines a first edge 111, and the intersection of the second leg 130 and the bottom panel 110 defines a second edge 113. The first leg 120 of the turbine shroud segment 100 may be shorter (i.e., extend in a radial direction over a shorter distance) than the second leg 130 of the shroud segment 100.
The first leg 120 and the second leg 130 may extend over a non-uniform length from the bottom panel 110. As illustrates in
Alternately, a single hole 127, 137 may be provided in the first leg 120 and the second leg 130. In this embodiment (not shown), the holes 127, 137 are aligned with one another, and mechanical fasteners positioned through the holes 127, 137 allow each shroud segment 100 to pivot onto its circumferentially adjacent shroud segment 100. Pressure loading ensures that the shroud segments 100 remain in contact, as described below, thus ensuring the desired sealing properties.
The turbine shroud segment 100, shown in closer detail in
As shown in
At the first end 150, the first leg 120 of the shroud segment 100 includes a first leg face 122, a first leg cutback face 124 spaced apart from the first leg face 122 in the circumferential direction, and a first leg intermediate surface 126 extending circumferentially between the first leg face 122 and the first leg cutback face 124. The second leg 130 of the shroud segment includes a second leg face 132, a second leg cutback face 134 spaced apart from the second leg face 132 in the circumferential direction, and a second leg intermediate surface 136 extending circumferentially between the second leg face 132 and the second leg cutback face 134.
The first leg face 122 and the second leg face 132 are integral with the first end face 152 of the bottom panel 110. The first leg cutback face 124 and the second leg cutback face 134 are integral with the first end cutback face 154 of the bottom panel 110. The first leg intermediate surface 126 and the second leg intermediate surface 136 are integral with the intermediate surface 156 of the bottom panel 110, thereby defining a U- or J-shaped channel. In the illustrated embodiment, the first leg intermediate surface 126 bisects the first leg 120 (that is, the first leg intermediate surface 126 is disposed at an equal distance in the axial direction between an outer surface of the first leg 120 and an inner surface of the first leg 120). Similarly, the second leg intermediate surface 136 bisects the second leg 130.
The second end face 162 extends radially inward from a radially inner (top) surface 117 of the bottom panel 110, and the second end cutback face 164 extends radially outward from the radially inner (bottom) surface 116 of the bottom panel 110. In the embodiment illustrated in
At the second end 160, the first leg 120 of the shroud segment 100 includes a third leg face 121, a third leg cutback face 123 spaced apart from the third leg face 121 in the circumferential direction, and a first leg projecting surface 125 extending circumferentially between the third leg face 121 and the third leg cutback face 123. The second leg 130 of the shroud segment includes a fourth leg face 131, a fourth leg cutback face 133 spaced apart from the fourth leg face 131 in the circumferential direction, and a second leg projecting surface 135 extending circumferentially between the fourth leg face 131 and the fourth leg cutback face 133.
The third leg face 121 and the fourth leg face 131 are integral with the second end face 162 of the bottom panel 110. The third leg cutback face 123 and the fourth leg cutback face 133 are integral with the second end cutback face 164 of the bottom panel 110. The first leg projecting surface 125 and the second leg projecting surface 135 are integral with the projecting surface 166 of the bottom panel 110, thereby defining a U- or J-shaped channel. In the illustrated embodiment, the first leg projecting surface 125 bisects the first leg 120 (that is, the first leg projecting surface 125 is disposed at an equal distance in the axial direction between an outer surface of the first leg 120 and an inner surface of the first leg 120). Similarly, the second leg projecting surface 135 bisects the second leg 130.
Although the end faces 152, 162; the end cutback faces 154, 165; the leg end faces 121, 122, 131, 132; and the leg cutback faces 123, 124, 133, 134 are illustrated as being parallel to a radial axis extending from the turbine rotor, it should be understood that faces may be oriented at some other angle, provided the mating surfaces of the female end 150 are angled in a complementary direction to the mating surfaces of the male end 160. An arrangement having one or more pairs of angled faces may further improve sealing properties by producing a more tortuous path for the hot combustion products.
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The present turbine shroud segments may be made of metal superalloys. However, for higher temperature environments, it may be desirable to make the shroud segments from ceramic matrix composite materials. When made of ceramic matrix composite material, the sealing surfaces (i.e., end faces and cutback faces of the bottom panel and the legs) may be finished to the desired tolerance by machining and/or grinding. Additionally, the ceramic matrix composite shroud segments may be coated, and the coating may be ground to achieve the desired sealing tolerance.
Exemplary embodiments of the present turbine shroud segments are described above in detail. The turbine shroud segments described herein are not limited to the specific embodiments described herein, but rather, components of the turbine shroud segments may be utilized independently and separately from other components described herein. For example, the turbine shroud segments described herein may have other applications not limited to practice with turbines for power-generating gas turbines, as described herein.
While the technical advancements have been described in terms of various specific embodiments, those skilled in the art will recognize that the technical advancements can be practiced with modification within the spirit and scope of the claims.
