BACKGROUND
The present disclosure relates generally to turbine engines and, more specifically, to systems and methods for use in controlling a clearance defined between rotating and stationary components.
At least some known gas turbine engines include a compressor, a combustor, and a high-pressure turbine coupled together in a serial flow relationship. More specifically, the compressor and high-pressure turbine are coupled through a shaft to form a high-pressure rotor assembly. Air entering the turbine engine is mixed with fuel and ignited to form a high energy gas stream. The high energy gas stream flows through the high-pressure turbine to rotatably drive the high-pressure turbine such that the shaft rotatably drives the compressor. In at least some known gas turbine engines, after being discharged from the high-pressure turbine, the gas stream continues to expand as it flows through a low-pressure turbine positioned aft of the high-pressure turbine. The low-pressure turbine includes a rotor assembly coupled to a drive shaft and a fan. The low-pressure turbine rotatably drives the fan through the drive shaft.
Known gas turbine engines operate at high temperatures to facilitate increasing engine performance and efficiency. However, operating at high temperatures can result in damage to hot gas path components if clearance gaps between the rotating and stationary components are not sufficiently large. Clearance is needed to enable the rotating components to rotate without rubbing against the stationary components. If the clearance is too large, combustion gases leak past the tips of the rotating components and do not efficiently drive the rotation of the turbine. If the clearance is too small, the rotating components rub against the stationary components and may cause vibration that may damage the turbine. Many know gas turbine engines are assembled with increased clearance gaps to facilitate the different rates of temperature change of the rotating and stationary components. However, increasing the gaps generally results in reducing the efficiency and performance of the gas turbine engine during startup and transient changes in operation of the gas turbine engine.
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BRIEF DESCRIPTION
In one aspect, a heat transfer system for a gas turbine engine is provided. The heat transfer system includes a first air duct extending axially within a wall of a single-shell casing of the gas turbine engine. The heat transfer system also includes a second air duct spaced radially from and extending substantially parallel to the first air duct. The second air duct is defined within the single-shell casing wall. Moreover, the heat transfer system includes a reversing chamber coupled in flow communication to the first air duct and the second air duct. The first air duct is configured to channel a fluid in a first axial direction relative to the gas turbine engine, and the second air duct is configured to channel the fluid in a second axial direction that is opposite the first axial direction.
In another aspect, a gas turbine engine defining a longitudinal axis is provided. The gas turbine engine includes a turbine and a single-shell casing having a wall housing the turbine. The gas turbine engine also has a heat transfer system formed in the single-shell casing. The heat transfer system includes a first air duct extending axially within the single-shell casing wall, and a second air duct spaced radially from and extending substantially parallel to the first air duct. The second air duct is defined within the single-shell casing wall. The first air duct is configured to channel a fluid in a first axial direction, and the second air duct is configured to channel the fluid in a second axial direction that is opposite the first axial direction.
In yet another aspect, a method for clearance control in a gas turbine engine is provided. The method includes channeling a fluid into a single-shell casing of the gas turbine engine. The method also includes directing the fluid into a first air duct extending axially within a wall of the single-shell casing in a first axial direction. Moreover, the method includes discharging the fluid into a reversing chamber coupled in flow communication with the first air duct and directing the fluid in a radially inward direction of the gas turbine engine. Furthermore, the method includes channeling the fluid into a second air duct extending
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substantially parallel to the first air duct and within the single-shell casing wall, and directing the fluid through the second air duct in a second axial direction that is opposite the first axial direction.
BRIEF DESCRIPTION OF THE DRAWINGS
These and other features, aspects, and advantages of the present disclosure will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:
FIG. 1 is a schematic illustration of an exemplary turbine engine assembly;
FIG. 2 is a sectional view of a portion of a high pressure turbine used with the turbine engine assembly shown in FIG. 1 and including a heat transfer system;
FIG. 3 is a perspective view of a section of a casing of the turbine engine assembly shown in FIG. 1;
FIG. 4 is an enlarged portion of the high pressure turbine shown in FIG. 2; and
FIG. 5 is a schematic side view of the heat transfer system shown in FIG. 2.
Unless otherwise indicated, the drawings provided herein are meant to illustrate features of embodiments of the disclosure. These features are believed to be applicable in a wide variety of systems comprising one or more embodiments of the disclosure. As such, the drawings are not meant to include all conventional features known by those of ordinary skill in the art to be required for the practice of the embodiments disclosed herein.
DETAILED DESCRIPTION
Unless otherwise indicated, approximating language, such as "generally," "substantially," and "about," as used herein indicates that the term so modified may apply to only an approximate degree, as would be recognized by one of ordinary skill in the art, rather than to an absolute or perfect degree. Approximating language may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to
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which it is related. Accordingly, a value modified by a term or terms, such as "about," "approximately," and "substantially," is not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations are identified. Such ranges may be combined and/or interchanged, and include all the sub-ranges contained therein unless context or language indicates otherwise.
As used herein, the terms "axial" and "axially" refer to directions and orientations that extend substantially parallel to a centerline of the turbine engine. Moreover, the terms "radial" and "radially" refer to directions and orientations that extend substantially perpendicular to the centerline of the turbine engine. In addition, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations that extend arcuately about the centerline of the turbine engine.
Additionally, unless otherwise indicated, the terms "first," "second," etc. are used herein merely as labels, and are not intended to impose ordinal, positional, or hierarchical requirements on the items to which these terms refer. Moreover, reference to, for example, a "second" item does not require or preclude the existence of, for example, a "first" or lower-numbered item or a "third" or higher-numbered item.
FIG. 1 is a schematic illustration of an exemplary turbine engine assembly 10. FIG. 2 is a cross-sectional view of a portion of high pressure turbine 20 of turbine engine assembly 10 (shown in FIG. 1). In the exemplary embodiment, turbine engine assembly 10 includes a gas turbine engine 12 that includes a low pressure compressor 14, a high pressure compressor 16, and a plurality of combustor assemblies 18 positioned downstream from high pressure compressor 16. Gas turbine engine 12 also includes a high pressure turbine 20 positioned downstream from combustor assemblies 18, a low pressure turbine 22 positioned downstream from high pressure turbine 20, and a power turbine 24 positioned downstream from low pressure turbine 22.
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In operation, a flow of intake air 26 is channeled through low pressure compressor 14 and a flow of compressor air (not shown) is channeled from low pressure compressor 14 to high pressure compressor 16. The compressor air is discharged from high pressure compressor 16 and channeled towards combustor assemblies 18 where the compressor air is mixed with fuel and combusted to form a flow of hot combustion gases 52 discharged towards high pressure turbine 20. The flow of hot combustion gases 52 discharged from combustor assembly 18 drives high pressure turbine 20 about a rotational axis 28 of gas turbine engine 12. The flow of combusted gas is channeled through low pressure turbine 22 and power turbine 24, where it is discharged from gas turbine engine 12 in the form of an exhaust gas flow 30.
Turbine engine assembly 10 also includes a heat transfer system 32 coupled in flow communication with high pressure compressor 16 to channel a flow of compressor discharge (CDC) air 41, extracted from a compressor discharge air port 42 of high pressure compressor 16, towards high pressure turbine 20 and low pressure turbine 22. As will be described in more detail below, CDC air 41 is channeled towards high pressure turbine 20 and low pressure turbine 22 to facilitate, for example, and without limitation, providing thermal management for a turbine casing 44 and hot gas path components, such as turbine blades 46 and nozzles 50, contained therein. As used herein, "hot gas path" refers to a flow path for hot combustion gases 52 within gas turbine engine 12, and "hot gas path component" refers to any component of gas turbine engine 12 that comes into contact with hot combustion gases 52 within hot gas path 54. In alternative embodiments, CDC air 41 is also used for other cooling and/or heating purposes. For example, and without limitation, CDC air 41 may be used to purge fluid within the wheel space of a rotor assembly.
In the exemplary embodiment, heat transfer system 32 incudes a control valve 34 and a controller 36 coupled in communication with control valve 34. Controller 36 is operable such that control valve 34 is selectively opened or closed, and when opened, either CDC air 41 or compressor stage air 38 is selectively channeled
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towards supply duct 40 via control valve 34. Compressor stage air 38 includes a flow of air that is extracted from at least one of a plurality of stages or extractions. For example, in one embodiment, air is extracted from extraction port 39 (not shown) of high pressure compressor 16, rather than from compressor discharge air port 42. In the exemplary embodiment, for example, and without limitation, air is extracted from compressor stage 8 (S8) or compressor stage 10 (S10). In some embodiments, controller 36 monitors a blade tip clearance gap "C" in high-pressure turbine 20, and is capable of selectively channeling either CDC air 41 and/or compressor stage air 38 through supply duct 40 based on clearance data, engine output data, and engine temperature data.
In the exemplary embodiment, gas turbine engine 12 includes annular casing 44 that houses a plurality of rows of turbine blades 46 that rotate about rotational axis 28. In the exemplary embodiment, turbine casing 44 is a single-shell casing. Within turbine casing 44, each row of turbine blades 46 is mounted to a turbine wheel 48. Between adjacent rows of turbine blades 46 are respective rows of nozzles 50 (broadly, guide vanes). Hot combustion gases 52 flow in hot gas path 54 through the rows of turbine blades 46 and nozzles 50.
Casing 44 facilitates forming an outer wall 56 of hot gas path 54, at least in part by securing one or more hot gas path components, such as turbine blades 46 and nozzles 50 thereto. An inner wall 58 of hot gas path 54 is defined near the outer rims of turbine wheels 48. In the exemplary embodiment, nozzles 50 are coupled to shroud components 60. Alternatively, nozzles 50 are coupled to turbine casing 44 at outer wall 56. In the exemplary embodiment, annular rows of shroud components 60 are coupled to turbine casing 44 and outer wall 58 and are aligned with blade tips 62 of turbine blades 46. Blade tip clearance gap "C" defined between shroud components 60 and blade tips 62 is referred to as the "clearance" or "clearance gap" of gas turbine engine 12.
In the exemplary embodiment, clearance gap "C" facilitates reducing an amount of hot combustion gases 52 that leak past blade tips 62. If clearance gap "C" is too small, blade tips 62 contact or scrape against shroud components 60, which
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causes wear to turbine blades 46 and shroud components 60, and can induce vibrations in gas turbine engine 12. Wear is generally undesirable as it facilitates increasing clearance gap "C" and can lead to damage to turbine blades 46 and shroud components 60. In addition, vibrations are generally undesirable because vibrations facilitate increasing a chance of damaging gas turbine engine 12.
FIG. 3 is a perspective view of a section of casing 44. FIG. 4 is an enlarged portion of casing 44. FIG. 5 is a schematic side view of heat transfer system 32 (shown in FIG. 2). In the exemplary embodiment, casing 44 is typically fabricated from a metallic material. An inner surface 64 of casing 44 includes a plurality of annular slots 66 sized and oriented to receive shroud components 60. Heat transfer system 32 includes an annular inlet cavity 96 that is integrally formed on an outer surface 98 of casing 44 and that is coupled in flow communication with supply duct 40. Inlet cavity 96 is oriented to receive a flow of compressed air (broadly a fluid), as indicated by serpentine arrow 80, from high pressure compressor 16. As described herein, compressed air 80 includes either CDC air 41 or compressor stage air 38 (shown in FIG. 2). Compressed air 80 flows into inlet cavity 96 where it is channeled to an inlet passage 72 and to a plurality of internal cooling passages, generally indicated at 70, all of which are part of heat transfer system 32. In the exemplary embodiment, annular plenum 76 includes a circumferential seal plate 100 that facilitates separating annular plenum 76 from an annular plenum 77.
Annular plenums 77 and 84 are formed in casing 44 and are sized and oriented to distribute compressed air to cooling passages 86 (shown in FIG. 2) formed in nozzles 50, and to various other cooling passages (not shown) formed in casing 44. As described herein, compressed air 80 is extracted from one or more stages (e.g., S8 or S10) or compressor discharge air port 42 of high pressure compressor 16. In one embodiment, annular plenum 84 formed around an upstream stage of nozzles 50 receives air compressed to a high pressure (therefore having an increased temperature) from a downstream stage of high pressure compressor 16, whereas annular plenum 77 surrounding a downstream stage of nozzles 50
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receives air compressed at a lower pressure from an upstream stage of high pressure compressor 16.
In the exemplary embodiment, cooling passages 70 extend longitudinally about rotational axis 28 of gas turbine engine 12. More specifically, cooling passages 70 follow a serpentine, e.g., a switch-back, path that reverses a direction of flow of compressed air 80 at a reversing chamber 88 defined near an axial end 90 of casing 44. Although only one set of cooling passages 70 is illustrated in FIG. 3, it is noted that several sets of cooling passages 70 are spaced circumferentially about casing 44. In the exemplary embodiment, reversing chamber 88 is formed in a forward end face 94 of casing 44, e.g., via a machining process, and is closed by a plate 92.
In operation, cooling passages 70 facilitate channeling compressed air 80 from annular plenum 76 through casing 44, and venting compressed air 80 into annular plenum 84. Heat transfer occurs between casing 44 and compressed air 80 as compressed air 80 flows through cooling passages 70. Compressed air 80 facilitates cooling casing 44 as casing 44 is at a higher temperature than compressed air 80. Typically, hot combustion gases 52 flowing through hot gas path 54 are hotter than compressed air 80. Alternatively, when compressed air 80 is at a higher temperature than casing 44, compressed air 80 facilitates increasing the temperature of casing 44. For example, compressed air 80 is typically hotter than casing 44 before combustion occurs in the combustor during startup of gas turbine engine 12.
In the exemplary embodiment, casing 44 is fabricated with predetermined thermal expansion characteristics. An amount of heat transfer between compressed air 80 and casing 44 is dependent on a surface area of cooling passages 70 and reversing chamber 88, which directly affects the thermal expansion and cooling of casing 44. As such, the thermal expansion characteristics of casing 44 are determined, in part, based on a length, shape, and cross-sectional area of cooling passages 70. In addition, a volume of cooling passages 70 and reversing chamber 88 directly affects a mass of casing 44, and the thermal expansion of casing 44. Accordingly,
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the thermal expansion characteristics of casing 44 may also be determined, in part, based on a volume of cooling passages 70.
In the exemplary embodiment, clearance gap "C" is determined, in part, based on the thermal expansion and contraction of casing 44 and the hot gas components, such as turbine blades 46 and turbine wheels 48. For example, during operation, casing 44 and the hot gas components, such as turbine blades 46 and turbine wheels 48, heat up during startup and steady-state operation of gas turbine engine 12 (or cool down during shutdown of gas turbine engine 12). Typically, however, casing 44, turbine blades 46, and turbine wheels 48, for example, heat up at different rates, causing clearance gap "C" to change size. In one embodiment, an amount of thermal expansion or contraction, in a radial direction, of turbine wheels 48 and turbine blades 46 is estimated to determine a radial displacement of blade tips 62. In addition, a radial displacement of casing 44 is estimated. A difference between the radial displacement of blade tips 62 and casing 44 determines clearance gap "C." Alternatively, controller 36 actively monitors clearance gap "C."
In the exemplary embodiment, cooling passages 70 includes a first air duct 102 coupled in flow communication to a second air duct 104 by reversing chamber 88, which is closed by plate 92 as described herein. More specifically, in the exemplary embodiment, first air duct 102 includes a first end portion 106 that extends to a second end portion 108 through an intermediate portion 110. First end portion 106 defines inlet 78, as described herein, that is in flow communication with annular plenum 76 while second end portion 108 is coupled in flow communication with reversing chamber 88. Second air duct 104 includes a first end portion 112 that extends from reversing chamber 88 to a second end portion 114 through an intermediate portion 116. Second end portion 114 defines outlet 82 and is coupled in flow communication to annular plenum 84. In an alternative embodiment, reversing chamber 88 includes an effusion plate 120 including a plurality of apertures 122 defined therethrough that facilitate generating a predetermined pressure drop between compressed air 80 exiting
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second end portion 108 of first air duct 102 and compressed air 80 entering first end portion 112 of second air duct 104.
First and second air ducts 102 and 104 extend generally axially though casing 44. In addition, first air duct 102 extends substantially parallel to second air duct 104 within casing 44 and generally radially outward from second air duct 104 at a similar general circumferential location in casing 44. The additional wall thickness of single-shell casing 44, as compared to double-shell casings, enables such a radial placement of first and second air ducts 102, 104. In addition, such an orientation of cooling passages 70 facilitates forming an increased number of circuits of cooling passages 70 in casing 44 than an arrangement including first and second air ducts 102, 104 positioned circumferential adjacent at the same radial location. Channeling compressed air 80 through first and second air ducts 102, 104 arranged in the manner described above facilitates reducing circumferential thermal gradients within casing 44. In addition, compressed air 80 passing within casing 44 facilitates reducing thermal gradients at annular slots 66 and shroud components 60, and thus, facilitates providing a desirable clearance benefit.
In the exemplary embodiment, first and second air ducts 102 and 104 are formed by a deep drilling process, such that first and second air ducts 102 and 104 each extend generally axially through casing 44. This is also facilitated by the additional wall thickness of single-shell casing 44. Deep drilling first and second air ducts 102, 104 facilitates ease of manufacturing and reduced fabrication cost as compared to double-shell casing turbine engines, which may require multiple operations to form such a cooling circuit.
In operation, compressed air 80 is channeled from high pressure compressor 16 (shown in FIG. 1) through supply duct 40 into inlet cavity 96. Compressed air 80 then flows through inlet 74 of inlet passage 72 and enters annular plenum 76. Compressed air 80 is then directed into inlet 78, along first air duct 102, and into reversing chamber 88. Compressed air 80 is channeled across plate surface 118 and into second air duct 104 before passing into and providing cooling for nozzles
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50 via annular plenum 84. Passing compressed air 80 through first air duct 102 in a first direction and through second air duct 104 in a second, opposite, direction generates a counter-flow within casing 44. In the exemplary embodiment, the counter-flow of compressed air 80 facilitates reducing circumferential thermal gradients within casing 44 by providing heat transfer between compressed air 80 flowing through first air duct 102 and second air duct 104.
The amount of compressed air 80 passing into annular plenum 76 and, more specifically, into heat transfer system 32 is utilized to control clearance gap "C" defined between blade tips 62 of turbine blades 46 and respective shroud components 60. More specifically, during startup of gas turbine engine 12, a clearance gap "C" defined between blade tips 62 of turbine blades 46 and respective shroud components 60 is larger than when gas turbine engine 12 is running at steady-state speed and at full steady-state load. Between startup and steady-state operations and between steady-state operation and shutdown, rotating components, i.e., turbine blades 46 and turbine wheels 48, of gas turbine engine 12 thermally expand and contract at a rate that is different than an expansion or contraction rate of stationary components such as casing 44 and shroud component 60. The different rates of thermal expansion or contraction lead to undesirable clearance gaps "C" defined between the rotating and stationary components. Controlling the flow of compressed air 80 into heat transfer system 32 facilitates controlling the thermal expansion and contraction rates of the rotating components and the stationary components while gas turbine engine 12 transitions between startup and steady-state operations. Controlling the expansion and contraction rates of the rotating components and the stationary components facilitates providing decreased clearance gaps "C" during transient and steady-state operations of gas turbine engine 12. The decreased clearance gaps "C" lead to a reduction in working fluid losses flowing past blade tip 62 of turbine blades 46, and thus facilitates improving gas turbine engine 12 performance and efficiency.
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Exemplary embodiments of a heat transfer system for a single-shell casing are described above in detail. The embodiments provide a counter-flow heat transfer system that provides advantages over know gas turbine engines in that, when the gas turbine engine is operating, the embodiments facilitate reducing bulk metal temperature and thermal gradients within a turbine portion of the gas turbine engine. The heat transfer system also provides deep convection cooling/heating to stationary components, such as the casing and shroud components, located along a hot gas path of the gas turbine engine. The counter-flow heat transfer system is arranged in a radial orientation that enables an increased number of cooling circuits formed in the casing as compared to known gas turbine engines. In addition, the heat transfer system is fabricated from a deep drilling process to facilitate reducing manufacturing costs. The heat transfer system facilitates controlling the thermal expansion and contraction of stationary turbine components and rotating turbine components. Moreover, the cooling flow may be selectively controlled to facilitate additional control of the thermal expansion and contraction rates of the stationary components and the rotating components through various operating states of the turbine, which facilitates reducing clearance gaps between the stationary components and the rotating components, particularly during the transition from one operating state to another operating state. The reduction in clearance gaps leads to a reduction in losses in working fluid along the hot gas path, facilitating improving performance and efficiency.
While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions, or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.
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The heat transfer systems and methods described above are not limited to the specific embodiments described herein, but rather, components of the apparatus and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein. For example, the exemplary embodiments can be implemented and utilized in connection with many other rotary machines.
Although specific features of various embodiments of the disclosure may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the disclosure, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing.
This written description uses examples to disclose the embodiments, including the best mode, and also to enable any person skilled in the art to practice the embodiments, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have 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.
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