The invention relates in general to gas turbine engines and, more specifically, to a vane carrier assembly for use in a gas turbine engine.
A conventional combustible gas turbine engine includes a compressor, a combustor, and a gas turbine. The engine further comprises an outer casing which defines an outer section for each of the compressor, combustor and gas turbine. A rotor extends through the engine. The rotor portion extending through the compressor is defined by a plurality of discs. Each disc can host a row of rotating airfoils, commonly referred to as blades. The rows of blades alternate with rows of stationary airfoils or vanes. The vanes can be mounted to the casing via one or more vane carrier assemblies. A clearance is defined between tips of the blades and an inner surface of vane carrier support panels. During operation of the gas turbine engine, fluid leakage through this clearance contributes to system losses, decreasing the operational efficiency of the engine. It is desirable to keep the clearance as small as possible to increase engine performance. However, it is necessary to maintain a clearance between the rotating and stationary components to prevent rubbing between the rotating and stationary components, which can lead to component or engine damage.
The size of the clearance can change during engine operation due to differences in the thermal growth response times of the compressor moving parts and that of the stationary structure. For example, the thermal growth response time of the stationary structure (e.g., the vane carrier assembly to which the vanes are connected) is significantly quicker than that of the rotating structure (rotor). Thus, the stationary structure has a faster thermal response time and responds (through expansion or contraction) more quickly to a change in temperature than the rotating structure.
In accordance with a first aspect of the present invention, a vane carrier assembly is provided for supporting vanes within a main engine casing of a gas turbine engine. The vane carrier assembly may comprise a plurality of vane support panels positioned adjacent to one another so as to define a vane support assembly. The support panels may be assembled such that the support panels expand circumferentially to minimize radial expansion of the vane support assembly during operation of the gas turbine engine. The vane carrier assembly may also comprise a control ring coupled to the main engine casing. The vane support assembly is coupled to the control ring.
The control ring may be supported by the main engine casing such that the main engine casing is capable of moving radially relative to the control ring.
The plurality of vane support panels may be made from a first material and the control ring may be made from a second material. The second material may be thermally more stable than the first material.
The first material may have a coefficient of thermal expansion greater than that of the second material.
The first material may be formed from a steel alloy.
The second material may be formed from one of INCOLOY® Alloy 909 (a nickel-iron-cobalt alloy), INCOLOY® Alloy 939 (a nickel based alloy), or NILO® Alloy K (a nickel-iron-cobalt controlled-expansion alloy).
In accordance with one embodiment, each of the vane support panels may comprise a first section, to which vanes are coupled, and a second section. The panel second sections may be formed from a first material and the panel first sections and the control ring may be formed from a second material. The first material may have a coefficient of thermal expansion greater than that of the second material.
In accordance with another embodiment, the control ring may have a radial dimension which is greater than an axial dimension.
In accordance with a further embodiment, the control ring may have an axial dimension which is greater than a radial dimension.
Each of the vane support panels may extend generally circumferentially in response to thermal expansion and contraction during operation of the gas turbine engine.
The control ring may be formed from a low thermal coefficient of expansion material to generally minimize thermal expansion and contraction of the control ring in a radial direction.
In accordance with a second aspect of the present invention, a method is provided for controlling clearance between tips of rotating blades and an inner surface of a vane support assembly within an engine casing of a gas turbine engine. The method may comprise providing a plurality of vane support panels positioned adjacent to one another to define the vane support assembly. The panels may be made of a first material. The method may further comprise providing a control ring adapted to be supported by the engine casing and made of a second material, and securing the vane support assembly to the control ring. The second material is thermally more stable than the first material.
In accordance with a third aspect of the present invention, a vane carrier assembly is provided for supporting vanes within a main engine casing of a gas turbine engine. The vane carrier assembly comprises a vane support assembly, and a control ring loosely coupled, axially supported and radially free in the illustrated embodiment, to the main engine casing such that the main engine casing is capable of moving radially relative to the control ring. The vane support assembly is coupled to the control ring.
While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:
In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, specific preferred embodiments in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
A gas turbine engine is provided comprising a compressor 10, a combustor (not shown) and a gas turbine (not shown). The gas turbine engine further comprises an outer casing 14, which defines an outer section for each of the compressor 10, the combustor and the gas turbine.
The compressor 10 comprises a plurality of rotor discs 20, which form part of a main engine rotor extending through the compressor 10, the combustor and the gas turbine. Each rotor disc 20 supports a row of rotating blades 22, which function to compress ambient air, which compressed air is provided to the combustor. The rows of blades 22 alternate with rows of stationary vanes. The vanes are mounted to the outer casing 14 via one or more vane carrier assemblies.
Compressor efficiency depends on tip clearance between the compressor rotor blades 22 and an inner surface of the one or more vane carrier assemblies. During operation of the gas turbine engine, fluid leakage through the clearance 322 between tips 22A of the rotor blades 22 and the inner surface of the one or more vane carrier assemblies contributes to system losses, decreasing the operational efficiency of the gas turbine engine. Hence, it is desirable to keep the clearance 322 as small as possible. However, it is necessary to maintain a clearance 322 between the rotating and stationary components during engine operation to prevent contact, such as rubbing, between the rotating and stationary components, which can lead to component damage, performance degradation, and extended downtime.
In accordance with a first embodiment of the present invention, a vane carrier assembly 40 is provided for supporting vanes 310A-310C within the engine outer casing 14, see
In the embodiment illustrated in
The vane support panels 42 are sized, shaped and assembled such that when they are at ambient temperature, e.g., from about 65 degrees to about 85 degrees, and an engine metal operating temperature, e.g., from about 750 degrees to about 850 degrees, edges 42A of adjacent vane support panels 42 never contact one another. Hence, as the panels 42 increase from ambient temperature to steady state operating temperature during startup and steady state operation of the gas turbine engine, the panel edges 42A do not engage one another. Because the panel edges 42A do not contact one another, the panels 42 are free to expand circumferentially as they increase in temperature. Since the panels 42 are free to expand circumferentially, they expand very little in a radial direction as they increase in temperature. As the panels 42 expand or move zero or very little in a radial direction, the panels 42 cause little or no radial movement of the vanes 310A-310C during start up, steady-state operation or cool down of the gas turbine engine.
A seal arrangement 50 is associated with the vane support panels 42 to prevent compressed gases from passing between adjacent edges 42A of the panels 42. In the illustrated embodiment, a seal 60, such as a conventional feather seal, extends axially along at least one axially extending side edge 42A of each panel 42, see
The vane support panels 42 defining the annular vane support assembly 44 are coupled to the control ring 41 via a sliding dovetail joint or firtree joint. The control ring 41 may be defined by two 180 degree control ring segments 41D and 41E, see
In conventional gas turbine engine compressors, typically all of the components of the engine casing, one or more vane carrier assemblies and the main engine rotor are made from a steel alloy material or other material having a high coefficient of thermal expansion. Further, the one or more vane carrier assemblies have a relatively low mass as compared to the main engine rotor. Because the one or more vane carrier assemblies and the main engine rotor are made from a material having a high coefficient of thermal expansion and the one or more vane carrier assemblies have a relatively low mass as compared to the rotor, the one or more vane carrier assemblies respond (through expansion or contraction) more quickly to a change in temperature than the rotor. Hence, the inner surfaces of the one or more vane carrier assemblies may move a radial distance at a greater rate than the rotor during engine start up and cool down. When the engine is stopped, the rotor, because of its large mass, cools down at a much slower rate than the vane carrier assemblies. Hence, once the engine is restarted after being stopped briefly following continuous engine operation, the rotor may be at an elevated temperature, while the vane carrier assemblies are cool. When the engine is restarted, the blade tips expand radially very quickly due to centrifugal forces before the vane carrier assemblies fully expand radially away from the blade tips. The clearance between the vane carrier assemblies and the blade tips must be sufficient to prevent contact when the rotor is at an elevated temperature and, hence, in a radially outwardly expanded condition, the blade tips are radially expanded due to centrifugal forces, and the vane carrier assemblies have not yet expanded radially away from the blade tips. So as to prevent contact between the blade tips and the vane carrier assemblies during an engine restart with the rotor hot, the initial build or cold clearance must be designed sufficiently large to prevent contact between the blade tips and the vane carrier assemblies.
As noted above, in the present invention, the control ring 41 is formed from a material having a coefficient of thermal expansion lower than that of the outer casing 14 and the main engine rotor. Further, the control ring flange 41B is loosely received in the annular recess 14A provided in the engine casing 14. Because the control ring flange 41B is loosely received in the annular recess 14A of the engine casing 14, the engine casing 14 is capable of moving radially relative to the flange 41B of the control ring 41 during engine startup, steady state operation and cool down without causing substantial radial movement of the control ring 41. Further, because the panel edges 42A do not engage one another as the panels 42 expand when heated during engine start up and steady-state operation, the panels 42 expand very little in a radial direction. It is noted that the control ring 41 may expand or contract radially a small amount when its temperature changes, causing a small amount of radial movement of the panels 42. Accordingly, the control ring 41 and the panels 42 move radially very little during engine startup, steady-state operation and cool down. Hence, the inner surfaces 242 of the panels 42 move very little radially relative to initial position of the blade tips 22A. Accordingly, it is believed that the clearance 322 between the inner surfaces 242 of the vane support panels 42 and the blade tips 22A varies by a smaller amount during engine startup, steady state operation and cool down in the present invention as compared to prior art gas turbine engines.
In the illustrated embodiment, a plurality of axially extending support beams 208 are coupled at first ends 208A to the control ring 41 via bolts 316 (shown in
In the illustrated embodiment, the vane carrier assembly 40 supports the last three rows of vanes 310A-310C of the compressor. It is contemplated that a vane carrier assembly constructed in accordance with the present invention could be used to support any row of vanes in the compressor or in the turbine where thermal response poses a performance debit.
It is contemplated that the vane support panels may be made from a material having a low coefficient of thermal expansion. However, typically such materials are more expensive than materials having higher coefficients of thermal expansion. Because the vane support panels 42 are sized, shaped and assembled so as to expand mainly in the circumferential direction and very little in the radial direction, and the panels 42 are relatively thin in the radial direction, it may be desirable to form the panels 42 from a material having a higher coefficient of thermal expansion as compared to the material from which the control ring 41 is formed so as to reduce costs. Accordingly, the amount of costly, low coefficient of thermal expansion material necessary to better control the clearance 322 between the blade tips 22A and the inner surfaces 242 of the vane support panels 42 is reduced.
It is noted that the control ring 41 of the vane carrier assembly in
A vane carrier assembly 400 constructed in accordance with a second embodiment of the present invention is illustrated in
A vane carrier assembly 500 constructed in accordance with a third embodiment of the present invention is illustrated in
Referring now to
When the conventional gas turbine is initially started, see plot 802, the blades expand outwardly in the radial direction very quickly to close the clearance. Soon thereafter, the vane carrier assembly expands radially outerwardly as it increases in temperature to increase the clearance. From about 50 time units to about 150 time units, the rotor starts to expand radially outerwardly as it increases in temperature to close the clearance. Steady state operation occurs from about 150 time units to about 1800 time units. At about 1800 time units, the engine trips, i.e., slows down. Because the blades are rotating slowly, the blade tips moved radially away from the vane carrier assembly, see the spike in the clearance, which occurs between about 1800-1850 time units. From about 1850 time units to about 1900 time units, the vane carrier assembly cools causing the clearance to reduce. The vane carrier assembly cools more rapidly than the rotor so there is initially a faster close down rate, followed by a slower close down rate. At about 2100 time units, the engine is restarted. Because the rotor is still at an elevated temperature, i.e., still radially expanded, the vane carrier assembly is cool and has not yet moved radially away from the blades and the blades quickly expand due to centrifugal forces, the clearance is nearly zero, see point 802A. Points 802A and 802B are minimum tip clearances, also called “pinch points,” for the conventional gas turbine engine. The “build” or “cold” clearance is designed to equal the difference between the steady state clearance and the pinch point 802A closest to zero, see Delta 806. The difference between these two values, i.e., Delta 806, is the desired build clearance to ensure the engine will not rub in operation.
When the gas turbine including the vane carrier assembly of the first embodiment of the present invention is initially started, the blades expand outwardly in the radial direction very quickly from about 0 time units to about 25 time units, see plot 804. From about 25 time units to about 150 time units, the control ring and the rotor expand radially outerwardly as they increases in temperature causing the clearance to increase. Steady state operation occurs from about 150 time units to about 1840 time units. At about 1840 time units, the engine trips, i.e., slows down. Because the blades are rotating slowly, the blade tips moved radially away from the vane carrier assembly, see the spike in the clearance at about 1840 time units. The control ring cools from about 1840 time units to about 1900 time units causing the vane support assembly to move toward the blade tips. Thereafter, the rotor begins to cool slightly moving the blade tips away from the vane support assembly. At about 2100 time units, the engine is restarted. Because the rotor is still at an elevated temperature, i.e., still radially expanded and the blades quickly expand due to centrifugal forces, the clearance is nearly zero, see point 804B. Points 804A, 804B and 802C are minimum tip clearances or “pinch points” for the gas turbine engine including the first embodiment design, with pinch point 804B being the one closest to zero. The “build” or “cold” clearance is designed to equal the difference between the steady state clearance and the pinch point having the lowest value, which is point 804B, see Delta 808. The difference between these two values, i.e., Delta 808, is the desired build clearance to ensure the engine will not rub in operation.
As is clear from plots 802 and 804, Delta 808 is less than Delta 806. Also, the steady state clearance for the vane carrier assembly of the present invention is less than the steady state clearance for the vane carrier assembly of the conventional engine, thereby increasing the efficiency of the compressor having the vane carrier assembly of the present invention.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.
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