Patent Publication Number: US-9835171-B2

Title: Vane carrier assembly

Description:
FIELD OF THE INVENTION 
     The invention relates in general to gas turbine engines and, more specifically, to a vane carrier assembly for use in a gas turbine engine. 
     BACKGROUND OF THE INVENTION 
     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. 
     SUMMARY OF THE INVENTION 
     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. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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: 
         FIG. 1  is a cross-sectional view of a vane carrier assembly constructed in accordance with a first embodiment of the present invention; 
         FIG. 2  is a perspective view of the vane carrier assembly of  FIG. 1 ; 
         FIG. 3  is a perspective view of a vane support panel according to the present invention; 
         FIG. 4  is a cross-sectional view of a vane carrier assembly constructed in accordance with a second embodiment of the present invention; 
         FIG. 5  is a cross-sectional view of a vane carrier assembly constructed in accordance with a third embodiment of the present invention; and 
         FIG. 6  are plots of measured clearances between blade tips and inner surfaces of vane support panels as a function of time for a conventional vane carrier assembly, and a vane carrier assembly constructed in accordance with the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     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  22 A 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  310 A- 310 C within the engine outer casing  14 , see  FIGS. 1-2 . The vane carrier assembly  40  comprises an annular control ring  41  and a plurality of vane support panels  42  coupled to the control ring  41 . The control ring  41  comprises a main body  41 A, a radial outer flange  41 B and a radial inner coupling member  41 C. The flange  41 B is loosely received in an annular recess  14 A provided in the engine casing  14 . In accordance with 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. For example, the outer casing  14  may be formed from a conventional steel alloy such as 2.25 Cr—Mo steel, the components of the main engine rotor may be formed from a conventional steel alloy, such as a NiCrMo alloy, while the control ring  41  may be formed from a material, such as 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), having a low thermal coefficient of expansion. During gas turbine engine startup, steady-state operation and cool-down, the outer casing  14  expands and contracts radially a greater amount than the control ring  41 . For example, during engine startup, the outer casing  14  expands radially a greater amount than the control ring. However, because the control ring flange  41 B is loosely received in the annular recess  14 A of the engine casing  14 , i.e., the flange  41 B is axially supported and radially free in the annular recess  14 A in the illustrated embodiment, the engine casing  14  is capable of moving radially relative to the flange  41 B of the control ring  41  during engine startup, steady state operation and cool down without causing substantial radial movement of the control ring  41 . Hence, thermally induced radial movement of the engine casing  14  has little influence on the radial location/position of the control ring  41 , i.e., the control ring  41  isn&#39;t caused to move a significant amount by the engine casing  14  in the radial direction during engine startup, steady state operation or cool down. The structure  14 B defining the annular recess  14 A functions to limit movement of the control ring flange  41 B and hence the control ring  41  in an axial direction, designed by arrow A in  FIG. 1 . The radial direction is designated by arrow R in  FIG. 1 . 
     In the embodiment illustrated in  FIG. 2 , ten substantially identical vane support panels  42  are provided. However, the number of panels  42  may be less than ten or greater than  10 . The plurality of vane support panels  42  are preferably formed in arcs, such that each may extend circumferentially through an angle falling within a range of from about 30 to about 60 degrees. The vane support panels  42  are assembled to define an annular vane support assembly  44 . In the illustrated embodiment, first, second and third rows of first, second and third vanes  310 A,  310 B and  310 C, respectively, are coupled to the panels  42  defining the vane support assembly  44 , see  FIG. 1 . The vanes  310 A- 310 C are coupled to the panels  42  via a sliding dovetail or firtree joint such that an extending member  311  of each vane  310 A- 310 C is received in a corresponding groove  142 A provided in a corresponding panel  42 , see  FIG. 1 . 
     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  42 A 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  42 A do not engage one another. Because the panel edges  42 A 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  310 A- 310 C 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  42 A 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  42 A of each panel  42 , see  FIGS. 1 and 4 . A seal  60  on one panel  42  extends across a gap between the one panel  42  and an adjacent panel  42  and is received in a groove of the adjacent panel  42  so as to prevent gases from passing through the gap between the edges  42 A during operation of the engine. The seals  60  do not prevent circumferential movement/expansion of the vane support panels  42 . It is also contemplated that a seal extending radially on at least one side edge  42 A of a panel  42  may be provided to prevent axially moving gases from moving through the gap between adjacent panels  42 . 
     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  41 D and  41 E, see  FIG. 2 . Prior to installation of the two control ring segments  41 D and  41 E into the engine casing  14 , panels  42 , e.g., five panels  42  in the illustrated embodiment, may be slidingly joined or connected to each control ring segment  41 D and  41 E. In the illustrated embodiment, an extending member or male portion  141 C of the control ring radial inner coupling member  41 C is received in a corresponding groove or female portion  142 B provided in each panel  42 . Once all panels  42  are coupled to the control ring segments  41 D and  41 E, the control ring segments  41 D and  41 E may be installed within the engine casing  14 . The vanes  310 A- 310 C are preferably coupled to the vane support panels  42  before the panels  42  are coupled to the control ring segments  41 D and  41 E. Bolts  312 A and  312 B extending through the control ring radial inner coupling member  41 C into the vane support panels  42  are provided for securing the vane support panels  42  to the control ring  41 . In the illustrated embodiment, a first bolt  312 A and a second bolt  312 B are centered on a corresponding panel  42  and are generally aligned in an axial direction. 
     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  41 B is loosely received in the annular recess  14 A provided in the engine casing  14 . Because the control ring flange  41 B is loosely received in the annular recess  14 A of the engine casing  14 , the engine casing  14  is capable of moving radially relative to the flange  41 B 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  42 A 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  22 A. Accordingly, it is believed that the clearance  322  between the inner surfaces  242  of the vane support panels  42  and the blade tips  22 A 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  208 A to the control ring  41  via bolts  316  (shown in  FIG. 1  but not shown in  FIG. 2 ). An annular main engine rotor cover  210  (shown in  FIG. 1  but not shown in  FIG. 2 ) is coupled via bolts  318  to second ends  208 B of the support beams  208 . Because the control ring  41  does not move radially or moves radially very little within the engine casing  14 , radial movement of the annular cover  210  is reduced making its position more stable within the gas turbine engine. 
     In the illustrated embodiment, the vane carrier assembly  40  supports the last three rows of vanes  310 A- 310 C 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  22 A 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  FIG. 1  has a radial dimension which is greater than an axial dimension. 
     A vane carrier assembly  400  constructed in accordance with a second embodiment of the present invention is illustrated in  FIG. 4 , where elements in the assembly  400  similar to those used in the assembly  40  illustrated in  FIG. 1  are referenced by the same numerals as used in  FIG. 1 . In this embodiment, the control ring  410  has an axial dimension which is greater than its radial dimension. The control ring  410  is smaller than the control ring  41  in the  FIG. 1  embodiment; hence, less low coefficient of thermal expansion material is required to form the control ring  410  of the  FIG. 4  embodiment as compared to the control ring  41  of the  FIG. 1  embodiment. The control ring  410  includes a flange  410 B that is loosely received in an annular recess  140 A provided in the engine casing  140 . Because the control ring flange  410 B is loosely received in the annular recess  140 A of the engine casing  140 , the engine casing  140  is capable of moving radially relative to the flange  410 B of the control ring  410  during engine startup, steady state operation and cool down without causing substantial radial movement of the control ring  410 . 
     A vane carrier assembly  500  constructed in accordance with a third embodiment of the present invention is illustrated in  FIG. 5 , where elements in the assembly  500  similar to those used in the assembly  40  illustrated in  FIG. 1  are referenced by the same numerals as used in  FIG. 1 . In this embodiment, each of the vane support panels  540  includes a first section  504  and a second section  506 . First, second and third rows of first, second and third vanes  310 A- 310 C, respectively, are coupled to the vane support panel first sections  504 , see  FIG. 5 . The first sections  504  are formed integral with a control ring  541  from a material, which preferably comprises a low coefficient of thermal expansion, such as 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). The control ring  541  includes a flange  541 A that is loosely received in an annular recess  14 A provided in the engine casing  14 . Because the control ring flange  541 A is loosely received in the annular recess  14 A of the engine casing  14 , the engine casing  14  is capable of moving radially relative to the flange  541 A of the control ring  541  during engine startup, steady state operation and cool down without causing substantial radial movement of the control ring  510 . The second section  506  of each vane support panel  540  is coupled to a corresponding first section  504  via one or more bolts  508  in the illustrated embodiment. The vane support panel second sections  506  may be formed from a material having a higher coefficient of thermal expansion than the material used to form the integral control ring  510  and vane support panel first sections  504  so as to reduce costs. 
     Referring now to  FIG. 6 , running tip clearances (i.e., units of distance along the Y axis) between blade tips and inner surfaces of vane support panels as a function of time (i.e., units of time along the X axis) for a conventional vane carrier assembly, corresponding to plot  802 , and a vane carrier assembly constructed in accordance with the embodiment illustrated in  FIG. 1 , corresponding to plot  804  have been compared. Zero on the Y axis (i.e., labeled as “Running Clearances”) implies that the clearance is zero, i.e., the blade tips and vane support panel inner surfaces are just touching. Clearance values above 0 on the Y axis are positive and correspond to an actual spacing between the blade tips and the inner surfaces of the vane support panels and clearance values below 0 on the Y axis are negative and correspond to engagement of the blade tips with the inner surfaces of the vane support panels.  FIG. 6  shows three different turbine operating cycles, a cold startup (from about 0 to about 1800 time units), a restart after a brief slowdown following continuous operation at steady state temperatures (from about 1800 to about 3800 time units) and a cool down (after about 3800 time units). 
     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  802 A. Points  802 A and  802 B 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  802 A 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  804 B. Points  804 A,  804 B and  802 C are minimum tip clearances or “pinch points” for the gas turbine engine including the first embodiment design, with pinch point  804 B 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  804 B, 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.