Patent Publication Number: US-2022213802-A1

Title: System for controlling blade clearances within a gas turbine engine

Description:
FIELD 
     The present disclosure generally pertains to gas turbine engines and, more particularly, to a system for controlling blade clearances within a gas turbine engine. 
     BACKGROUND 
     A gas turbine engine generally includes a compressor section, a combustion section, and a turbine section. During operation, the compressor section progressively increases the pressure of air entering the engine and supplies this compressed air to the combustion section. The compressed air and a fuel mix within the combustion section and burn within a combustion chamber to generate high-pressure and high-temperature combustion gases. The combustion gases flow through a hot gas path defined by the turbine section before exiting the engine. In this respect, the turbine section converts energy from the combustion gases into rotational energy. Specifically, the turbine section includes a plurality of rotor blades, which extract kinetic energy and/or thermal energy from the combustion gases flowing therethrough. The extracted rotational energy is, in turn, used to rotate one or more shafts, thereby driving the compressor section and/or a fan assembly of the gas turbine engine 
     In general, it desirable to minimize the clearance between the outer tips of the rotor blades and the adjacent shrouds or drum to maximize the amount of energy extracted by the rotor blades. However, the rotor blades expand and contract relative to the shrouds/drum during thermal cycling of the engine. As such, the clearance between the rotor blades and the shrouds/drum generally decreases as the engine heats up. In this respect, when the clearance between the blade tips and the shrouds/drum is minimized during cold operation of the engine, the blade tips may contact the shrouds/drum when the engine heats up. Conversely, when the clearance between the blade tips and the shroud/drum is optimized for hot operation, such clearance may be sufficiently large to reduce the efficiency of the energy extraction during cold operation of the engine. 
     Accordingly, an improved system for controlling blade clearances within a gas turbine engine would be welcomed in the technology. 
     BRIEF DESCRIPTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     In one aspect, the present subject matter is directed to a system for controlling blade clearances within a gas turbine engine. The gas turbine engine defines an axial centerline and a radial direction extending orthogonal to the axial centerline. The system includes a rotor disk and a rotor blade coupled to the rotor disk. Additionally, the system includes an outer turbine component positioned outward of the rotor blade in the radial direction such that a clearance is defined between the rotor blade and the outer turbine component. Furthermore, the system includes a heat exchanger configured to receive a flow of cooling air bled from the gas turbine engine and transfer heat from the received flow of the cooling air to a flow of coolant to generate cooled cooling air. Moreover, the system includes a valve configured to control the flow of the coolant to the heat exchanger. In this respect, the cooled cooling air is supplied to at least one of the rotor disk or the rotor blade to adjust the clearance between the rotor blade and the outer turbine component. 
     In another aspect, the present subject matter is directed to a system for controlling blade tip clearances within a gas turbine engine. The gas turbine engine defines an axial centerline and a radial direction extending orthogonal to the axial centerline. The system includes an inner rotor configured to rotate in a first direction and an inner rotor blade coupled to the inner rotor. Additionally, the system includes an outer rotating drum configured to rotate in a second direction opposite of the first direction and an outer rotor blade coupled to the outer rotating drum. Furthermore, the system includes heat exchanger configured to receive a flow of cooling air bled from the gas turbine engine and transfer heat from the received flow of the cooling air to a flow of coolant to generate cooled cooling air. In addition, the system includes a first air valve configured to direct a first portion of the cooled cooling air to the outer rotating drum and a second portion of the cooled cooling air to cool the inner rotor and a second air valve configured to direct a first portion of the cooling air to the outer rotating drum and a second portion of the cooling air to cool the inner rotor. As such, the cooled cooling air is supplied to at least one of the outer rotating drum or the inner rotor to adjust a first clearance defined between the inner rotor blade and the outer rotating drum and a second clearance between the outer rotor blade and the inner rotor. 
     These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG. 1  is a schematic cross-sectional view of one embodiment of a gas turbine engine of an aircraft; 
         FIG. 2  is a schematic cross-sectional view of another embodiment of a gas turbine engine of an aircraft; 
         FIG. 3  is a schematic view of one embodiment of a system for controlling blade clearances within a gas turbine engine; 
         FIG. 4  is an enlarged, partial schematic view of the system for controlling blade clearances within a gas turbine engine shown in  FIG. 3 , particularly illustrating a rotor disk and a rotor blade of the gas turbine engine; 
         FIG. 5  is a cross-sectional side view of one embodiment of a turbine section of a gas turbine engine; 
         FIG. 6  is a schematic view of another embodiment of a system for controlling blade clearances within a gas turbine engine; and 
         FIG. 7  is another schematic view of the system shown in  FIG. 6 . 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION 
     Reference now will be made in detail to exemplary embodiments of the presently disclosed subject matter, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation and should not be interpreted as limiting the present disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present disclosure without departing from the scope or spirit of the present disclosure. For instance, features illustrated or described as part of one embodiment can be used with 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. 
     As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. 
     Furthermore, the terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     Additionally, the terms “low,” “high,” or their respective comparative degrees (e.g., lower, higher, where applicable) each refer to relative parameter magnitudes (e.g., speeds, pressures, or temperatures) within an engine, unless otherwise specified. For example, a “low-pressure turbine” operates at a pressure generally lower than a “high-pressure turbine.” Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low-pressure turbine” may refer to the lowest maximum pressure turbine within a turbine section, and a “high-pressure turbine” may refer to the highest maximum pressure turbine within the turbine section. 
     In general, the present subject matter is directed to a system for controlling blade clearances within a gas turbine engine. As will be described below, the gas turbine engine includes a shaft, a rotor disk coupled to the shaft, and a rotor blade coupled to the rotor disk (e.g., via a dovetail connection) such that the rotor blade extends outward from the disk along a radial direction of the engine. Additionally, the gas turbine engine includes an outer turbine component, such as a shroud or a counter-rotating outer drum, positioned outward of the rotor blade in the radial direction. As such, a clearance is defined between the outer tip of the rotor blade and the outer turbine component. 
     The disclosed system is configured to supply cooled cooling air to the rotor disk and/or the rotor blade to adjust the clearance between the rotor blade and the outer turbine component. Specifically, the system includes a heat exchanger configured to receive a flow of cooling air bled from the gas turbine engine. For example, in one embodiment, the cooling air is bled from a compressor discharge plenum of the engine. As such, the heat exchanger is configured to transfer heat from the received flow of cooling air to a flow of coolant (e.g., supercritical carbon dioxide) to generate cooled cooling air. Additionally, the system includes a valve configured to control the flow of the coolant to the heat exchanger to adjust the temperature of the cooled cooling air. The cooled cooling air is, in turn, routed to the rotor disk and/or the rotor blade to adjust the clearance between the rotor blade and the outer turbine component. For example, in some embodiments, the cooled cooling air flows from the heat exchanger to the rotor disk and/or the rotor blade through a conduit at least partially positioned between the shaft and a combustor of the engine. 
     The disclosed system provides one or more technical advantages. For example, as described above, the disclosed system supplies cooled cooling air to the rotor disk and/or the rotor blade. Such cooled cooling air reduces the amount that rotor blade expands as the engine heats up, thereby controlling clearance between the rotor blade and outer turbine component. Furthermore, as mentioned above, the temperature of the cooled cooling air may be controlled by the valve. In this respect, increasing the amount of and/or decreasing the temperature of the cooled cooling air supplied to the rotor blade and/or the rotor disk may shrink the rotor blade and/or the disk, thereby increasing the clearance via a reduction in the blade tip radius. Conversely, the clearance may be decreased by reducing the amount of and/or increasing the temperature of cooling air supplied to the rotor blade and/or the rotor disk. Moreover, the disclosed system allows the thermal expansion/contraction of the rotor blade and/or the disk to be controlled independently of the thermal expansion/contraction of the outer turbine component. 
     Referring now to the drawings,  FIG. 1  is a schematic cross-sectional view of one embodiment of a gas turbine engine  10 . In the illustrated embodiment, the engine  10  is configured as a high-bypass turbofan engine. However, in alternative embodiments, the engine  10  may be configured as a propfan engine, a turbojet engine, a turboprop engine, a turboshaft gas turbine engine, or any other suitable type of gas turbine engine. 
     As shown in  FIG. 1 , the engine  10  defines a longitudinal direction L, a radial direction R, and a circumferential direction C. In general, the longitudinal direction L extends parallel to an axial centerline  12  of the engine  10 , the radial direction R extends orthogonally outward from the axial centerline  12 , and the circumferential direction C extends generally concentrically around the axial centerline  12 . 
     In general, the engine  10  includes a fan  14 , a low-pressure (LP) spool  16 , and a high pressure (HP) spool  18  at least partially encased by an annular nacelle  20 . More specifically, the fan  14  may include a fan rotor  22  and a plurality of fan blades  24  (one is shown) coupled to the fan rotor  22 . In this respect, the fan blades  24  are spaced apart from each other along the circumferential direction C and extend outward from the fan rotor  22  along the radial direction R. Moreover, the LP and HP spools  16 ,  18  are positioned downstream from the fan  14  along the axial centerline  12  (i.e., in the longitudinal direction L). As shown, the LP spool  16  is rotatably coupled to the fan rotor  22 , thereby permitting the LP spool  16  to rotate the fan  14 . Additionally, a plurality of outlet guide vanes or struts  26  spaced apart from each other in the circumferential direction C extend between an outer casing  28  surrounding the LP and HP spools  16 ,  18  and the nacelle  20  along the radial direction R. As such, the struts  26  support the nacelle  20  relative to the outer casing  28  such that the outer casing  28  and the nacelle  20  define a bypass airflow passage  30  positioned therebetween. 
     The outer casing  28  generally surrounds or encases, in serial flow order, a compressor section  32 , a combustion section  34 , a turbine section  36 , and an exhaust section  38 . For example, in some embodiments, the compressor section  32  may include a low-pressure (LP) compressor  40  of the LP spool  16  and a high-pressure (HP) compressor  42  of the HP spool  18  positioned downstream from the LP compressor  40  along the axial centerline  12 . Each compressor  40 ,  42  may, in turn, include one or more rows of stator vanes  44  interdigitated with one or more rows of compressor rotor blades  46 . Moreover, in some embodiments, the turbine section  36  includes a high-pressure (HP) turbine  48  of the HP spool  18  and a low-pressure (LP) turbine  50  of the LP spool  16  positioned downstream from the HP turbine  48  along the axial centerline  12 . Each turbine  48 ,  50  may, in turn, include one or more rows of stator vanes  52  interdigitated with one or more rows of turbine rotor blades  54 . 
     Additionally, the LP spool  16  includes the low-pressure (LP) shaft  56  and the HP spool  18  includes a high pressure (HP) shaft  58  positioned concentrically around the LP shaft  56 . In such embodiments, the HP shaft  58  rotatably couples the rotor blades  54  of the HP turbine  48  and the rotor blades  46  of the HP compressor  42  such that rotation of the HP turbine rotor blades  54  rotatably drives HP compressor rotor blades  46 . As shown, the LP shaft  56  is directly coupled to the rotor blades  54  of the LP turbine  50  and the rotor blades  46  of the LP compressor  40 . Furthermore, the LP shaft  56  is coupled to the fan  14  via a gearbox  60 . In this respect, the rotation of the LP turbine rotor blades  54  rotatably drives the LP compressor rotor blades  46  and the fan blades  24 . 
     In several embodiments, the engine  10  may generate thrust to propel an aircraft. More specifically, during operation, air (indicated by arrow  62 ) enters an inlet portion  64  of the engine  10 . The fan  14  supplies a first portion (indicated by arrow  66 ) of the air  62  to the bypass airflow passage  30  and a second portion (indicated by arrow  68 ) of the air  62  to the compressor section  32 . The second portion  68  of the air  62  first flows through the LP compressor  40  in which the rotor blades  46  therein progressively compress the second portion  68  of the air  62 . Next, the second portion  68  of the air  62  flows through the HP compressor  42  in which the rotor blades  46  therein continue progressively compressing the second portion  68  of the air  62 . The compressed second portion  68  of the air  62  is subsequently delivered to the combustion section  34 . In the combustion section  34 , the second portion  68  of the air  62  mixes with fuel and burns to generate high-temperature and high-pressure combustion gases  70 . Thereafter, the combustion gases  70  flow through the HP turbine  48  which the HP turbine rotor blades  54  extract a first portion of kinetic and/or thermal energy therefrom. This energy extraction rotates the HP shaft  58 , thereby driving the HP compressor  42 . The combustion gases  70  then flow through the LP turbine  50  in which the LP turbine rotor blades  54  extract a second portion of kinetic and/or thermal energy therefrom. This energy extraction rotates the LP shaft  56 , thereby driving the LP compressor  40  and the fan  14  via the gearbox  60 . The combustion gases  70  then exit the engine  10  through the exhaust section  38 . 
       FIG. 2  is a schematic cross-sectional view of another embodiment of a gas turbine engine  10  of an aircraft. Like the embodiment of the engine  10  shown in  FIG. 1 , the embodiment of the engine  10  shown in  FIG. 2  includes an LP turbine  50 . However, unlike the embodiment of the engine  10  shown in  FIG. 1 , in the embodiment of the engine  10  shown in  FIG. 2 , the LP turbine  50  is a counter-rotating turbine. Specifically, in such an embodiment, the LP turbine  50  includes an inner rotor  72  configured to rotate in a first direction (e.g., one of the clockwise or counter-clockwise directions) and one or more rows of inner rotor blades  74  coupled to and extending outward from the inner rotor  72  in the radial direction R. Furthermore, in such an embodiment, the LP turbine  50  includes an outer rotating drum  76  configured to rotate in a second direction opposite of the first direction (e.g., the other of clockwise or counter-clockwise directions) and one or more rows of outer rotor blades  78  extending inward from the drum  102  toward the axial centerline  12  in the radial direction R. As shown, the rows of outer rotor blades  78  are interdigitated with the rows of inner rotor blades  74 . In addition, the LP shaft  24  may be coupled to the outer rotor  76  of the LP turbine  50  via a gearbox  80 . 
     The configuration of the gas turbine engine  10  described above and shown in  FIG. 1  is provided only to place the present subject matter in an exemplary field of use. Thus, the present subject matter may be readily adaptable to any manner of gas turbine engine configuration, including other types of aviation-based gas turbine engines, marine-based gas turbine engines, and/or land-based/industrial gas turbine engines. 
       FIG. 3  illustrates one embodiment of a system  100  for controlling blade clearances within a gas turbine engine. In general, the system  100  will be discussed in the context of the gas turbine engine  10  described above and shown in  FIGS. 1 and 2 . However, the disclosed system  100  may be implemented within any gas turbine engine having any other suitable configuration. 
     As shown, in several embodiments, the combustion section  34  of the gas turbine engine  10  includes one or more combustors  102 . In general, the combustor(s)  102  is positioned outward from the shafts  56 ,  58  along the radial direction R Each combustor  102  includes a liner  104  defining a combustion chamber  106  therein. Moreover, each combustor  102  includes one or more fuel nozzles  108 , which supply a mixture of fuel and compressed air (e.g., the compressed, the second portion  68  of the air  62 ) to the combustion chamber  106 . The fuel and air mixture burns within the combustion chamber  106  to generate the high-temperature and high-pressure combustion gases  70 . Although  FIG. 3  illustrates a single combustor  102 , the combustion section  34  may, in other embodiments, include a plurality of combustors  102 . 
     Additionally, in several embodiments, the combustion section  34  includes a compressor discharge casing  110 . In such embodiments, the compressor discharge casing  110  at least partially surrounds or otherwise encloses the combustor(s)  102  in the circumferential direction C. In this respect, a compressor discharge plenum  112  is defined between the compressor discharge casing  110  and the liner  104 . The compressor discharge plenum  112  is, in turn, configured to supply compressed air to the combustor(s)  102 . Specifically, as shown, the compressed air exiting the HP compressor  42  is directed into the compressor discharge plenum  112  by an inlet guide vane  113 . The compressed air within the compressor discharge plenum  112  will be referred to as compressed air  114 . A portion of the compressed air  114  is supplied to the combustion chamber(s)  106  of the combustor(s)  102  by the fuel nozzle(s)  108  for use in combusting the fuel. As will be described below, in some embodiments, another portion of the compressed air  114  is used for cooling components of the HP turbine  48  of the gas turbine engine  10 . 
     As shown, the system  100  includes a heat exchanger  116 . More specifically, the heat exchanger  116  is configured to receive a flow of cooling air (indicated by arrow  118 ) bled from the gas turbine engine  10  and a flow of coolant (indicated by arrows  120 ). In this respect, the heat exchanger  116  is configured to transfer heat from the flow of the cooling air  118  to the flow of coolant  120 . Such heat transfer cools the received cooling air  118 , thereby generating cooled cooling air (indicated by arrows  122 ). As will be described below, the temperature of the cooled cooling air  122  may be adjusted by controlling the volume of the coolant  120  flowing through the heat exchanger  116 . Thereafter, the cooled cooling air  122  is routed to the turbine section  36  to control the blade tip clearances therein. 
     In several embodiments, the heat exchanger  116  is configured to receive the cooling air  118  from the compressor discharge plenum  112 . Specifically, in such embodiments, a portion of the compressed air  114  is bled from the compressor discharge plenum  112  and routed to the heat exchanger  116 . For example, in one embodiment, the system  100  includes a conduit  124  that conveys the compressed air  114  from the compressor discharge plenum  112  to the heat exchanger  116 . Although not shown in  FIG. 3 , a suitable valve(s) may be provided in associated with the conduit  124  to control the flow of the compressed air  114  from the compressor discharge plenum  112  to the heat exchanger  116 . However, in alternative embodiments, the cooling air  118  received by the heat exchanger  116  may be bled from any other suitable location on the gas turbine engine  10 , such as the compressor section  32 . 
     The heat exchanger  116  may be positioned at any suitable location within the gas turbine engine  10 . For example, as shown, in one embodiment, the heat exchanger  116  is positioned outward along the radial direction R from the combustor(s)  102 . 
     Additionally, the flow of coolant  120  received by the heat exchanger  116  may be formed from any suitable type of coolant. For example, in one embodiment, the flow of coolant  120  may be a flow of supercritical carbon dioxide. 
     As mentioned above, in some embodiments, the temperature of the cooled cooling air  122  may be adjusted by controlling the flow of the coolant  120  to the heat exchanger  116 . In such embodiments, the system  100  includes a valve  126  configured to control the flow of the coolant  120  to the heat exchanger  116  and a bypass conduit  128 . More specifically, the valve  126  is configured to adjust the volume of the coolant  120  supplied to the heat exchanger  116  by allowing a portion of the coolant to bypass the heat exchanger  116  via the bypass conduit  128 . For example, the valve  126  may increase the volume of the coolant  120  supplied to the heat exchanger  116  by allowing less coolant  120  (or no coolant  120 ) to flow into the bypass conduit  128 . Such an increase in the volume of the coolant  120  supplied to the heat exchanger  116  decreases the temperature of the cooled cooling air  122 . Conversely, the valve  126  may decrease the volume of the coolant  120  supplied to the heat exchanger  116  by allowing more coolant  120  to flow into the bypass conduit  128 . Such a decrease in the volume of the coolant  120  supplied to the heat exchanger  116  increases the temperature of the cooled cooling air  122 . 
     Referring now to  FIGS. 3 and 4 , the cooled cooling air  122  is routed to the turbine section  36  to control the blade tip clearances therein. In several embodiments, the cooled cooling air  122  may be used to control the blade tip clearances of a first stage  130  of the HP turbine  48 . However, in alternative embodiments, the cooled cooling air  122  may be used to control the blade tip clearances of any other blade tips within the turbine section  36 . 
     In general, the first stage  130  includes a row of circumferentially arranged stator vanes  52  (one is shown) and a row of circumferentially arranged rotor blades  54  (one is shown). As shown, the stator vanes  52  are positioned downstream from the combustion chamber  106  relative to the direction of the flow of the combustion gases  70 . As such, the stator vanes  52  define a downstream end of the compressor discharge plenum  112 . Furthermore, the rotor blades  54  are positioned downstream from the stator vanes  52  in the direction of the flow of the combustion gases  70 . In this respect, the stator vanes  52  and rotor blades  52  partially form a hot gas path  132  along which the combustion gases  70  flow through the turbine section  36 . More specifically, each stator vane  52  includes inner and outer bands  134 ,  136  respectively forming the inner and outer boundaries of the hot gas path  132  in the radial direction R. Each stator vane  54  also includes an airfoil  138  extending through the hot gas path  132  along the radial direction R between the inner and outer bands  134 ,  136 . Moreover, each rotor blade  54  includes a base portion  140  and an airfoil  142  extending outward in the radial direction R from the base portion  140  into the hot gas path  132 . The base portion  140  of each rotor blade  54  is coupled to a rotor disk  144  (e.g., via a dovetail connection, a fir tree-type connection, etc.), with the rotor disk  144 , in turn, being coupled to the HP shaft  58 . As such, rotation of the rotor disk  144  and the rotor blades  54  rotate the HP shaft  58 , which, in turn, drives the compressor  32  as described above. 
     Moreover, in some embodiments, one or more seals may be positioned adjacent of the rotor disk  144 . For example, as shown in  FIG. 4 , inner and outer seals  143 ,  145  are positioned upstream of the rotor disk  144  along the axial centerline  12  relative to the direction of the flow of the combustion gases  70  through the gas turbine engine  10 . In such an embodiment, the inner seal  143  is positioned inward along the radial direction R of the outer seal  145  such that a gap  147  is defined therebetween. As will be described below, the cooled cooling air  122  may flow through the gap  147  toward the rotor disk  144  and then outward along the radial direction R between the outer seal  145  and the rotor disk  144 , thereby cooling the rotor disk  144  and the rotor blade  54 . 
     Additionally, the first stage  130  of the HP turbine  48  includes one or more outer turbine components  146  partially defining the hot gas path  132 . In general, the outer turbine component(s)  146  is positioned outward of airfoil  142  of the rotor blade  54  in the radial direction R such that the component(s)  146  define an outer boundary of the hot gas path  132  in the radial direction R. As shown, a clearance (indicated by arrow  148 ) is defined between the tips  150  of the airfoils  142  of the rotor blades  54  and an inner radial surface(s)  152  of the outer turbine component(s)  146 . As will be described below, the clearance  148  may be controlled by the cooled cooling air  122  supplied to the first stage  130 . For example, in the illustrated embodiment, the outer turbine component(s)  146  is a shroud  154  enclosing the rotor blades  54 . However, in alternative embodiments, the outer turbine component(s)  146  may be any other suitable component(s), such as a counter-rotating drum (e.g., the outer rotating drum  76 ) or a shroud attached to a counter-rotating drum. 
     Furthermore, in several embodiments, the system  100  includes a conduit  156 . In general, the conduit  156  is configured to supply the cooled cooling air  122  from the heat exchanger  116  to the rotor blades  54  and the rotor disk  144  of the first stage  130 . As such, in some embodiments, the conduit  156  is at least partially positioned between the combustor(s)  102  (and, more specifically, the inner portion of the compressor discharge casing  110  in the radial direction R) and the HP shaft  58  in the radial direction R. Additionally, in some embodiments, the system  100  includes an inducer  157  configured to direct the cooled cooling air flowing through the conduit  156  toward the rotor disk  144 . For example, as shown, in one embodiment, the inducer  157  narrows as the inducer  157  extends from the downstream end of the conduit  156  toward the rotor disk  144  to direct the cooled cooling air  122  through the gap  147 . 
     The conduit  156  may have any suitable configuration for routing the cooled cooling air  122  to the rotor disk  144  and/or the rotor blade  54 . For example, as mentioned above, in the illustrated embodiment, the heat exchanger  116  is positioned outward from the combustor(s)  102  in the radial direction R. In such an embodiment, the conduit includes a first portion  158  extending along the radial direction R from the heat exchanger  116  inward toward the axial centerline  12 . In one embodiment, the first portion  158  of the conduit  156  is positioned upstream of the combustor(s)  102  relative to the direction of flow of the combustion gases  70  through the gas turbine engine  10 . In addition, in several embodiments, the system  100  includes a valve  159  configured to control the flow of the cooled cooling air  122  from the heat exchanger  116  to the cooling passage  156 . Furthermore, the conduit  156  includes a second portion  161  extending from the downstream end of the first portion  158  along the axial centerline  12  between the HP shaft  58  and the combustor  102  toward the rotor disk  144 . As indicated above, the inducer  157  is positioned at the downstream end of the second portion  161  to direct the cooled cooling air  121  exiting the conduit  156  through the gap  147  and toward the rotor disk  144 . However, in alternative embodiments, the conduit  156  may have any other suitable configuration. 
     In some embodiments, the flow of cooled cooling air  122  supplied to the rotor disk  144  and/or the rotor blade  54  by the conduit  156  is supplemented with additional compressed air  114  from the compressor discharge plenum  112 . More specifically, as shown in  FIG. 4 , in such embodiments, the inner radial side of the compressor discharge casing  110  defines a bleed port  160  fluidly coupling the compressor discharge plenum  112  and the cooling passage  156 . In this respect, a portion of the compressed air  114  from the compressor discharge plenum  112  flows through the bleed port  160  and directly into the cooling passage  156 . This additional compressed air  114  may increase the volume of the cooling air  122  supplied to the turbine section  36 , thereby increasing the cooling capacity of such air  122  without increasing the size of the heat exchanger  116 . In one embodiment, a valve (not shown) may control the flow the additional compressed air  114  through the bleed port  160 . 
     Referring particularly to  FIG. 4 , in several embodiments, the cooled cooling air  122  flowing through the conduit  156  is supplied to the rotor disk  144  and the rotor blades  54  of the first stage  130  of the HP turbine  48 . More specifically, the cooled cooling air  122  flows inward along the radial direction R from the heat exchanger  116  through the first portion  158  of the conduit  156  and subsequently downstream relative to the direction of flow of the combustion gases  70  through the second portion  161  of the conduit  156 . The inducer  157  then directs the cooled cooling air  122  exiting the conduit  156  through the gap  147  between the seals  143 ,  145  and onto the rotor disk  144  of the first stage  130 . The cooled cooling air  122  then flows outward in the radial direction R between the outer seal  145  and a forward or upstream surface  162  of the rotor disk  144  such that the cooled cooling air  122  cools the disk  144 . Thereafter, the cooled cooling air  122  flows along forward or upstream surfaces  164  of the the base portions  140  of the first stage rotor blades  54 . In one embodiment, a portion of the cooled cooling air  122  flows through passages  166  (one is shown) defined by the base portions  140  of the first stage rotor blades  54 , thereby cooling the interiors of the rotor blades  54 . 
     As indicated above, the cooled cooling air  122  allows the clearance  148  between the rotor blade tips  150  and the outer turbine component(s)  146  to be controlled. More specifically, the cooling of the first stage rotor disk  144  and rotor blades  54  provided by the cooled cooling air  122  causes the disk  144  and the rotor blades  54  to shrink in the radial direction R. In this respect, increasing the amount of and/or decreasing the temperature (e.g., by controlling the valve  126 ) of the cooled cooling air  122  supplied to the rotor disk  144  and the rotor blades  54  increases the amount such components shrink, thereby increasing the clearance  152 . Conversely, decreasing the amount of and/or increasing the temperature (e.g., by controlling the valve  126 ) of the cooled cooling air  122  supplied to the rotor disk  144  and the rotor blades  54  causes the components to grow, thereby decreasing the clearance  152 . As such, the disclosed system  100  allows the clearance  148  to be minimized as the temperature of the gas turbine engine  10  varies during operation. 
     After cooling the first stage rotor disk  144  and rotor blades  54 , the spent cooled cooling air  122  may be exhausted into the hot gas path  132 . For example, in some embodiments, at least a portion of the spent cooled cooling air  122  may flow along the upstream surfaces  164  of the rotor blades  54  and be exhausted in the hot gas path through a clearance  168 . The clearance  168  is, in turn, defined between the inner bands  134  of the stator vanes  52  and the platforms of the rotor blades  54 . Moreover, in some embodiments, at least a portion of the spent cooled cooling air  122  may flow through the passages  166  in the base portions  140  of the first stage rotor blades  54  and be exhausted in the hot gas path through an outlet  170 . However, in alternative embodiments, the spent cooled cooling air  122  may be exhausted into the hot gas path  132  in any other suitable manner. 
     In several embodiments, the first stage outer turbine component(s)  146  are cooled in a controlled manner to further control the size of the clearance  148  between the outer turbine component(s)  146  and the rotor blade tips  150 . Specifically, in such embodiments, compressed air  114  from the compressor discharge plenum  112  is supplied to the outer turbine component(s)  146  to cool this component(s)  146 , thereby shrinking the component(s)  146 . Shrinking the outer turbine component(s)  146 , in turn, decreases the clearance  148 . For example, in one embodiment, the shroud  154  (e.g., a 360-degree shroud) defines a passage  172  through which the compressed air  114  flows to the cool the shroud  154 . However, in other embodiments, the compressed air  114  may be simply directed at the outer radial side the outer turbine component(s)  146 . Furthermore, the compressed air  114  may be supplied to a turbine case  173  to which the outer turbine component(s)  146  is coupled to adjust the clearance between the rotor blade tip(s)  150  and the outer turbine component(s)  146 . After cooling the outer turbine component(s)  146 , the spent compressed air  114  may be exhausted into the hot gas path  132 . In other embodiments, the air supplied to cool the outer turbine components, such as a 360-degree ring shroud or a counter-rotating drum (with or without attached segmented shrouds), may be cooled cooling air cooled by an independent heat exchanger having a coolant (e.g., supercritical CO2) which is controlled by an independent valve. This cooled cooling air  122  used to cool the outer turbine components may also be controlled or metered using an in-line air valve, such as the valve  159 . 
     The flow of the cooled cooling air  122  to the first stage rotor disk  144  and the rotor blades  54  may be controlled independently of the flow of the compressed air  114  or cooled cooling air  122  to the outer turbine component(s)  146 . As such, the clearance  148  between the outer turbine component(s)  146  and rotor blade tips  150  may be adjusted by controlling the flow and temperature of the cooled cooling air  122  to the first stage rotor disk  144  and the rotor blades  54 , the flow of the compressed air  114  or the flow and temperature of the cooled cooling air  122  to the outer turbine component(s)  146 , or both. 
     Additionally, in some embodiments, the system  100  may be used to control the sizes of the clearances in counter-rotating turbines. More specifically, as shown in  FIG. 5 , in such a turbine (e.g., the LP turbine  50  shown in  FIG. 2 ), a first clearance (indicated by arrow  174 ) is defined between the tips  176  of the airfoils of the inner rotor blades  74  and an inner radial surface(s)  178  of the outer rotating drum  76 . Moreover, a second clearance (indicated by arrow  180 ) is defined between the tips  182  of the airfoils of the outer rotor blades  78  and an outer radial surface(s)  184  of the inner rotor  72 . 
       FIG. 6  is a schematic view of another embodiment of a system  100  for controlling blade clearances within a gas turbine engine. Like the embodiment of the system  100  shown in  FIGS. 3 and 4 , the system  100  shown in  FIG. 6  includes a heat exchanger  116  configured to receive and cool cooling air  118  to generate cooled cooling air  122 . However, unlike the embodiment of the system  100  shown in  FIGS. 3 and 4 , the system  100  shown in  FIG. 6  includes a first air valve  186  in fluid communication with the heat exchanger  116 . In this respect, the first air valve  186  is configured to direct or otherwise route a first portion  188  of the cooled cooling air  122  from the heat exchanger  116  to the outer rotating drum  76  and a second portion  190  of the cooled cooling air  122  from the heat exchanger  116  to cool the inner rotor  72 . Furthermore, unlike the embodiment of the system  100  shown in  FIGS. 3 and 4 , the system  100  shown in  FIG. 6  includes a second air valve  192  configured to route a first portion of cooling air  118  (e.g., cooling air  118  bled from the compressor discharge plenum  112 , but not delivered to the heat exchanger  116 ) to the outer rotating drum  76  and a second portion  196  of the cooling air  118  to cool the inner rotor  72 . 
     As indicated above, the cooling air  118  and the cooled cooling air  122  allow the first and second clearances  174  and  180  to be controlled. More specifically, the cooling of the inner rotor  72  and the outer rotating drum  76  provided by the cooling air  118  and the cooled cooling air  122  causes the inner rotor  72  and the outer rotating drum  76  to shrink in the radial direction R. In this respect, increasing the amount of cooled cooling air  122  and the decreasing the amount of cooled air  118  (e.g., by controlling the valves  186 ,  192 ) supplied to the inner rotor  72  and the outer rotating drum  76  increases the amount such components shrink, thereby increasing the clearance  174 ,  180 . Conversely, decreasing the amount of cooled cooling air  122  and the increasing the amount of cooled air  118  (e.g., by controlling the valves  186 ,  192 ) supplied to the inner rotor  72  and the outer rotating drum  76  causes the components to grow, thereby decreasing the clearance  152 . As such, the disclosed system  100  allows the clearances,  174 ,  180  to be minimized as the temperature of the gas turbine engine  10  varies during operation. 
     As shown in  FIGS. 6 and 7 , the cooled cooling air  122  and the cooling air  118  are delivered to the outer rotating drum  76  and the inner rotor  72  via angled nozzles  198  to further affect the cooling of the outer rotating drum  76  and the inner rotor  72 . More specifically, the first portion  188  of the cooled cooling air  122  may be introduced to or otherwise directed at the outer rotating drum  76  through one or more angled nozzles  198  such that a tangential component of the velocity of the first portion  188  of the cooled cooling air  122  is in the second direction (i.e., the direction in which the outer rotating drum  76  rotates). Furthermore, the second portion  190  of the cooled cooling air  122  may be introduced to or otherwise directed at the inner rotor  72  through one or more angled nozzles  198  such that a tangential component of the velocity of the second portion  190  of the cooled cooling air  122  is in the first direction (i.e., the direction in which the inner rotor  72  rotates). Directing the cooled cooling air  122  in the same direction as the rotation of the inner rotor  72  and the outer rotating drum  76  increases the cooling that the cooled cooling air  122  provides. Conversely, the first portion  194  of the cooling air  118  may be introduced to or otherwise directed at the outer rotating drum  76  through one or more angled nozzles  198  such that a tangential component of the velocity of the first portion  194  of the cooling air  118  is in the first direction (i.e., the opposite direction to which the outer rotating drum  76  rotates). Moreover, the second portion  196  of the cooling air  118  may be introduced to or otherwise directed at the inner rotor  72  through one or more angled nozzles  198  such that a tangential component of the velocity of the second portion  196  of the cooling air  118  is in the second direction (i.e., the opposite direction in which the inner rotor  72  rotates). Directing the cooled cooling air  122  in the opposite direction as the rotation of the inner rotor  72  and the outer rotating drum  76  decreases the cooling that the cooled cooling air  122  provides. In this respect, the controlling the amount of cooling air  118  and the cooled cooling air  122  (e.g., with the valves  186 ,  192 ) and its direction of flow relative to the inner rotor  72  and the outer rotating drum  76  (e.g., via the nozzles  198 ), the clearances,  174 ,  180  to be minimized as the temperature of the gas turbine engine  10  varies during operation. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims. 
     Further aspects of the invention are provided by the subject matter of the following clauses: 
     A system for controlling blade clearances within a gas turbine engine, the gas turbine engine defining an axial centerline and a radial direction extending orthogonal to the axial centerline, the system comprising: a rotor disk; a rotor blade coupled to the rotor disk; an outer turbine component positioned outward of the rotor blade in the radial direction such that a clearance is defined between the rotor blade and the outer turbine component; a heat exchanger configured to receive a flow of cooling air bled from the gas turbine engine and transfer heat from the received flow of the cooling air to a flow of coolant to generate cooled cooling air; and a valve configured to control the flow of the coolant to the heat exchanger, wherein the cooled cooling air is supplied to at least one of the rotor disk or the rotor blade to adjust the clearance between the rotor blade and the outer turbine component. 
     The system of one or more of these clauses, further comprising: a shaft coupled to the rotor disk such that rotation of the rotor disk and the rotor blade rotates the shaft; a combustor positioned outward in the radial direction from the shaft; and a conduit at least partially positioned between the shaft and the combustor in the radial direction such that the cooled cooling air flows through the cooling passage from the heat exchanger to the at least one of the rotor disk or the rotor blade. 
     The system of one or more of these clauses, further comprising: an inducer configured to direct the cooled cooling air flowing through the conduit toward the rotor disk. 
     The system of one or more of these clauses, wherein the inducer narrows as the inducer extends from the conduit toward the rotor disk. 
     The system of one or more of these clauses, further comprising: a seal positioned upstream of the rotor disk along the axial centerline relative to a direction of flow through the gas turbine engine, wherein the inducer directs the cooled cooling air such that the cooled cooling air flows between the rotor disk and the seal. 
     The system of one or more of these clauses, wherein the seal corresponds to an outer seal, the system further comprising: an inner seal positioned inward along the radial direction relative to the outer seal such that a gap is defined between the inner and outer seals through which the cooled cooling air flows from the inducer toward the rotor disk. 
     The system of one or more of these clauses wherein the conduit includes a first portion extending along the radial direction from the heat exchanger and a second portion extending along the axial centerline from the first portion toward the rotor disk. 
     The system of one or more of these clauses, wherein the first portion of the conduit is positioned upstream of the combustor relative to a direction of flow through the gas turbine engine. 
     The system of one or more of these clauses, wherein the heat exchanger is positioned outward along the radial direction from the combustor. 
     The system of one or more of these clauses, further comprising: a compressor discharge casing at least partially surrounding the combustor, the compressor discharge casing defining a compressor discharge plenum configured to supply compressed air to the combustor, wherein the cooling air received by the heat exchanger is bled from the compressor discharge plenum. 
     The system of one or more of these clauses, further comprising: a turbine case coupled to the outer turbine components, wherein the cooled cooling air is supplied to the turbine case to adjust the clearance between the rotor blade and the outer turbine component. 
     The system of one or more of these clauses, further comprising: a bypass conduit fluidly coupled to the valve such that the bypass conduit is configured to permit at least a portion of the coolant to bypass the heat exchanger. 
     The system of one or more of these clauses, wherein the cooled cooling air is discharged into a hot gas path at least partially defined by the rotor blade and the outer turbine component after being supplied to the at least one of the rotor disk or the rotor blade. 
     The system of one or more of these clauses, wherein the coolant comprises supercritical carbon dioxide. 
     The system of one or more of these clauses, wherein the outer turbine component comprises a shroud or an outer rotating drum. 
     A system for controlling blade tip clearances within a gas turbine engine, the gas turbine engine defining an axial centerline and a radial direction extending orthogonal to the axial centerline, the system comprising: an inner rotor configured to rotate in a first direction; an inner rotor blade coupled to the inner rotor; an outer rotating drum configured to rotate in a second direction opposite of the first direction; an outer rotor blade coupled to the outer rotating drum; a heat exchanger configured to receive a flow of cooling air bled from the gas turbine engine and transfer heat from the received flow of the cooling air to a flow of coolant to generate cooled cooling air; a first air valve configured to direct a first portion of the cooled cooling air to the outer rotating drum and a second portion of the cooled cooling air to cool the inner rotor; and a second air valve configured to direct a first portion of the cooling air to the outer rotating drum and a second portion of the cooling air to cool the inner rotor, wherein the cooled cooling air is supplied to at least one of the outer rotating drum or the inner rotor to adjust a first clearance defined between the inner rotor blade and the outer rotating drum and a second clearance between the outer rotor blade and the inner rotor. 
     The system of one or more of these clauses, where the first portion of the cooled cooling air is introduced to the outer rotating drum through an angled nozzle such that a tangential component of a velocity of the first portion of the cooled cooling air is in the second direction. 
     The system of one or more of these clauses, where the second portion of the cooled cooling air is introduced to the inner rotor through an angled nozzle such that a tangential component of a velocity of the second portion of the cooled cooling air is in the first direction. 
     The system of one or more of these clauses, where the first portion of the cooling air is introduced to outer rotating drum through an angled nozzle such that a tangential component of a velocity of the first portion of the cooling air is in the first direction. 
     The system of one or more of these clauses, where the second portion of the cooling air is introduced to inner rotor through an angled nozzle such that a tangential component of a velocity of the second portion of the cooling air is in the second direction.