Patent Publication Number: US-9840932-B2

Title: System and method for blade tip clearance control

Description:
BACKGROUND 
     The subject matter disclosed herein relates to a system and method for reducing blade tip clearances of turbomachines. In particular, the present disclosure relates to a system and method for reducing blade tip clearances by controlling axial displacement of turbomachine components. 
     Traditionally, turbomachines include a turbine with rotating blades within a stationary turbomachine shroud. A clearance may be included between a tip of each blade and the turbomachine shroud. This clearance may be referred to as a blade tip clearance. Blade tip clearances enable combustion gases passing through the turbomachine to leak over the tips of the blades, between the blade tips and the turbomachine shroud. Leakage of combustion gases in this manner may reduce an overall efficiency of the turbomachine system, and particularly the turbomachine itself. Thus, it is now recognized that there is a need for a system and method for improving, reducing, or eliminating blade tip clearances. 
     BRIEF DESCRIPTION 
     Certain embodiments commensurate in scope with the originally claimed invention are summarized below. These embodiments are not intended to limit the scope of the claimed invention, but rather these embodiments are intended only to provide a brief summary of possible forms of the invention. Indeed, the invention may encompass a variety of forms that may be similar to or different from the embodiments set forth below. 
     In a first embodiment, a system includes a turbomachine rotor having a shaft and turbomachine blades coupled to the shaft. The system also includes a turbomachine stator having a shroud surrounding the turbomachine blades of the turbomachine rotor. Further, the system includes a cooling channel having at least a first portion of the cooling channel extending upstream of a final stage of a compressor of the system, where the cooling channel is configured to receive cooled compressed air from the compressor and direct the cooled compressed air adjacent to the turbomachine rotor to reduce thermal expansion and/or axial displacement of the turbomachine rotor. 
     In a second embodiment, a method for reducing blade tip clearances of a turbomachine includes diverting a first portion of compressed air to a heat exchanger during certain stages of operation of the turbomachine and cooling the first portion of compressed air via the heat exchanger to generate a cooled compressed air. The method also includes routing the cooled compressed air through a channel proximate a rotor of the turbomachine, where the channel includes at least a first portion of the channel extending upstream of a final stage of a compressor of the turbomachine. Further, the method includes cooling the rotor to effectuate a reduction in thermal expansion and/or axial displacement of the rotor to reduce a blade tip clearance between a blade of the turbomachine and a stator of the turbomachine. 
     In a third embodiment, a system includes a turbomachine rotor having a shaft and turbomachine blades coupled to the shaft. The system also includes a turbomachine stator having a shroud surrounding the turbomachine blades of the turbomachine rotor. Further, the system includes a cooling channel having at least a first portion of the cooling channel extending upstream of a final stage of a compressor of the system, where the cooling channel is configured to receive cooled compressed air from the compressor and direct the cooled compressed air adjacent to the turbomachine rotor to reduce thermal expansion and/or axial displacement of the turbomachine rotor. The system also includes a control system. The control system is configured to selectively enable fluid communication between the compressor and the cooling channel. The control system includes a valve disposed between the compressor and the cooling channel, where the valve is configured to be selectively opened to enable fluid communication between the compressor and the cooling channel based on an operating condition or stage of operation of the turbomachine system. The control system also includes a sensor disposed proximate the cooling channel and configured to detect a parameter relating to the operating condition of the turbomachine system. Further, the control system includes a controller configured to receive the parameter relating to the operating condition or stage of operation of the turbomachine system and, based on the operating condition or stage of operation, selectively open or close the valve to enable fluid communication between the compressor and the cooling channel. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein: 
         FIG. 1  is a schematic diagram of an embodiment of a turbomachine system having an axial displacement control system, in accordance with aspects of the present disclosure; 
         FIG. 2  is a cross-sectional side view of an embodiment of a turbomachine blade and a honeycomb structure disposed on a turbine shroud, in accordance with aspects of the present disclosure; 
         FIG. 3  is a cross-sectional side view of an embodiment of the turbomachine blade and the honeycomb structure of  FIG. 2 , without a blade tip clearance, in accordance with aspects of the present disclosure; 
         FIG. 4  is a cross-sectional side view of an embodiment of a turbomachine system having an axial displacement control system, in accordance with aspects of the present disclosure; and 
         FIG. 5  is a process flow diagram of a method for controlling blade tip clearances, in accordance with aspects of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     One or more specific embodiments of the present invention will be described below. In an effort to provide a concise description of these embodiments, all features of an actual implementation may not be described in the specification. It should be appreciated that in the development of any such actual implementation, as in any engineering or design project, numerous implementation-specific decisions must be made to achieve the developers&#39; specific goals, such as compliance with system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time consuming, but would nevertheless be a routine undertaking of design, fabrication, and manufacture for those of ordinary skill having the benefit of this disclosure. 
     When introducing elements of various embodiments of the present invention, the articles “a,” “an,” “the,” and “said” are intended to mean that there are one or more of the elements. The terms “comprising,” “including,” and “having” are intended to be inclusive and mean that there may be additional elements other than the listed elements. 
     Embodiments of the present disclosure include a turbomachine (e.g., a turbomachine system) having a turbomachine stator and a turbomachine rotor. The turbomachine may include a compressor and/or a turbine, such as a gas turbine, a steam turbine, a hydro turbine, or any combination thereof. In the following discussion, embodiments of a clearance control system are discussed in context of a gas turbine, but are equally applicable to other types of turbines as well. 
     The stator of the turbomachine is stationary and may include a compressor shroud, compressor vanes, a turbine shroud, turbine vanes, and an optional transition shroud between the compressor shroud and turbine shroud. The rotor may include a shaft and compressor blades and turbine blades coupled to the shaft, where the rotor components rotate about a rotational axis extending through the shaft. A compressor of the turbomachine system includes the compressor shroud, compressor vanes of the stator, and the compressor blades of the rotor, while a turbine of the turbomachine system includes at least the turbine shroud and turbine vanes of the stator and the turbine blades of the rotor. The compressor blades and compressor vanes alternate in stages along the rotational axis, and the turbine blades and turbine vanes alternate in stages along the rotational axis. The shaft of the rotor extends through both the compressor and the turbine and, as previous described, is coupled to the compressor blades and turbine blades. Thus, as the shaft rotates, so too do the compressor blades and turbine blades, where each stage of the compressor blades and turbine blades are disposed between stages of the compressor vanes and turbine vanes, respectively. However, it should be noted that, in accordance with present embodiments, the turbine may be a multishaft turbine. For example, a separate shaft (e.g., a load shaft) may be coupled between the turbine and a load, such that rotation of the turbine blade rotates the load shaft to drive the load. Any number of shafts may be included in the turbine for rotating various components of the turbine. 
     The turbine blades may cut into or physically contact an adradable structure such as metallic honeycomb. The honeycomb structure may be disposed on the stationary turbine shroud while the turbine blades rotate with the shaft during operation. By contacting the honeycomb structure during operation (e.g., during rotation), the turbine blades block hot combustion gases being routed through the turbine from leaking over tips of the turbine blades between the turbine blades and the honeycomb structure disposed on the turbine shroud. However, due to thermal expansion of various components of the turbomachine, the turbine blades may axially separate from the honeycomb structure during various operating conditions or stages of operation (e.g., in an axial direction parallel to the rotational axis). The distance between the tip of each blade and the honeycomb structure of the stationary turbine shroud, while the turbine blade tip is separated from the honeycomb structure, may be referred to as a blade tip clearance. An axial blade tip clearance (e.g., longitudinal blade tip clearance) may refer to a blade tip clearance measured axially from the blade tip to the honeycomb structure, i.e., in the axial direction relative to the rotational axis. A radial blade tip clearance may refer to a blade tip clearance measured radially from the blade tip to the honeycomb structure, i.e., in a radial direction perpendicular to the rotational axis. 
     To reduce or eliminate blade tip clearances (in particular, axial blade tip clearances), embodiments of the present disclosure include an axial displacement control system, or control system for short. The axial displacement control system may control axial displacement of various turbomachine components of the stator and/or rotor at least in part by utilizing compressed air or a portion of compressed air generated by the compressor, or some other type of coolant, such as an inert gas (e.g., nitrogen), or any other gas, liquid, or vapor. In particular, the axial displacement control system may control axial displacement of portions of the rotor with respect to the stator. For example, a portion of the compressed air generated by the compressor may be exported to a heat exchanger for cooling. The portion of compressed air may then be cooled and routed proximate portions of the rotor for cooling the rotor. By cooling the rotor with the cooled compressed air, the axial displacement of the rotor may be reduced compared to embodiments where the rotor is not cooled with the cooled compressed air. In turn, by cooling the rotor, the axial displacement of the turbine blades, which are coupled to or, in other words, are a part of the rotor, is also reduced. By reducing axial displacement of the turbine blades, blade tips of the turbine blades may remain in contact with the honeycomb structure. In other words, by reducing axial displacement of the turbine blades, axial blade tip clearance may be reduced or eliminated. 
     The control system may also control axial displacement of other components of the turbomachine besides the rotor. For example, the control system may substantially divert the cooled compressed air only to the rotor, or mostly to the rotor, such that the stator is allowed to heat and expand. Thus, while the turbine blades of the rotor “contract” opposite the axial direction into the honeycomb structure disposed on the turbine shroud (or, more accurately, are blocked from expanding away from the honeycomb structure in the axial direction), the turbine shroud (of the stator) may thermally expand in the axial direction into the blades (of the rotor) to further facilitate closure of the axial blade tip clearance. Indeed, other mechanisms may be utilized for ensuring that the stator thermally expands more so than the rotor. For example, specific materials may be selected for turbine components proximate the area that is cooled by the cooled compressed air. Materials of rotor components may have a low coefficient of thermal expansion and materials of stator components may have a high coefficient of thermal expansion, at least relative to one another. For example, steel alloys with varying amounts of Iron, Aluminum, Boron, Carbon, Chromium, Cobalt, Copper, Lead, Manganese, Molybdenum, Nickel, Phosphorus, Silicon, Sulfur, Tantalum, Titanium, Thallium, Tungsten, and Zirconium may be used for components of the rotor and/or stator. Common names for such alloys include Stainless Steel, Inconel, and Chrom-Moly Alloys. By selecting appropriate materials, thermal expansion of the rotor in the axial direction may be reduced compared to thermal expansion of the stator in the axial direction, which may reduce blade tip clearances as set forth in the present disclosure. 
     By thermally expanding in the axial direction, the stator (or, more specifically, the honeycomb structure disposed on the stator) may be axially displaced into the tips of the turbine blades. By varying between, or simultaneously facilitating, (a) cooling of the rotor and (b) heating of the stator, the rotor and corresponding turbine blades are blocked from axial growth away from the stator and the stator axially expands into or toward the turbine blades of the rotor. The control system may, depending on operating conditions or stages of operation, determine if, when, and/or how much rotor cooling and/or stator heating (or simply less or no cooling) is appropriate or desirable. The control system and turbomachine components will be described in detail below with reference to the figures. 
     Turning now to the figures,  FIG. 1  is a schematic diagram of an embodiment of a turbomachine system  10  having a compressor  12 , combustors  14 , fuel nozzles  16 , and a turbine  18 . The fuel nozzles  16  route a liquid fuel and/or gas fuel, such as natural gas or syngas, into the combustors  14 . The combustors  14  also receive compressed air  19  generated by the compressor  12  for mixing with the fuel, and the combustors  14  ignite and combust the fuel-air mixture. Hot, pressurized combustion gases  20  (e.g., exhaust) are then passed from the combustors  14  into the turbine  18 . The turbine  18  includes turbine blades  22  and a turbine shroud  23 , where the turbine blades  22  are coupled to a rotary shaft  24 , and the turbine shroud  23  is stationary with respect to the shaft  24  and the turbine blades  22 . Coupled to the turbine shroud  23  are a number of turbine vanes  25 , which direct or alter flow (e.g., by controlling pressure/velocity of the flow) of the hot pressured combustion gases  20  between each set of turbine blades  22 . Thus, as the hot pressurized combustion gases  20  pass through the turbine  18 , the turbine blades  22  of the turbine  12  rotate and drive the shaft  24  into rotation, and the turbine vanes  25  prepare the hot pressured combustion gases  20  for each successive stage of turbine blades  22 . 
     The shaft  24  also extends through the compressor  12 , among other components of the system  10 , and rotates about a rotational axis  26  extending through the shaft  24 . The compressor  12  comprises a number of compressor blades  28  which are coupled to the shaft  24 . Thus, as the shaft  24  rotates via driven rotation of the turbine blades  22  as described above, the compressor blades  28  also rotate. The compressor  12  is configured to receive air (e.g., ambient air), and the air is compressed in the compressor  12  as the blades  28  of the compressor  12  rotate and as a cross-sectional area of the compressor  12  decreases in an axial direction  30  of the compressor  12  parallel to the rotational axis  26 . Similar to the turbine  18 , the compressor  12  also includes a compressor shroud  32 , which is stationary with respect to the shaft  24  and the compressor blades  26 . The compressor  12  likewise includes compressor vanes  34 , which may redirect or alter pressure/velocity of the flow of air through the compressor  12  as the air is compressed. The compressor vanes  34  may be coupled to the compressor shroud  32 , such that the compressor vanes  34  are stationary with respect to the rotating shaft  24  and components coupled to the shaft  24  (e.g., the compressor blades  28  and turbine blades  22 ). 
     Ultimately, the turbomachine system  10  may drive a load  36 , which may be coupled to the shaft  24  or to a separate shaft that is coupled to a final stage of the blades  22  of the turbine  18 . In other words, in some embodiments, some of the blades  22  of the turbine  18  may be used for driving the shaft  24 , the compressor  12 , and the turbine  18 , while others of the blades  22  may be used for driving a different shaft that drives the load  36 . In the illustrated embodiment, for clarity, the shaft  24  is coupled to all rotary components of the illustrated schematic gas turbine engine  10 , including the load  36 . 
     Often, the rotary or rotating components of the turbomachine system  10  are collectively referred to as a rotor. The rotor in the illustrated embodiment, for example, may include at least the shaft  24 , the compressor blades  28 , and the turbine blades  22 . Further, stationary components of the turbomachine system  10  are often referred to, collectively, as a stator. The stator in the illustrated embodiment, for example, may include at least the compressor shroud  32 , the compressor vanes  34 , the turbine shroud  23 , the turbine vanes  25 , and optionally a transition shroud  38  disposed between the compressor shroud  32  and the turbine shroud  23 . In some embodiments, the optional transition shroud  38  may be replaced with a rotating cover (which may be a part of the rotor) or may not be included at all. For example, in some embodiments, the compressor shroud  32  may seamlessly transition into the turbine shroud  23 , or the compressor shroud  32  and the turbine shroud  23  may be disposed proximate each other. 
     To enhance efficiency of the turbine  18 , clearance between the stationary turbine shroud  23  and tips of the turbine blades  22  may be reduced. This clearance may be referred to as a blade tip clearance. Blade tip clearance may actually include two components: axial blade tip clearance and radial blade tip clearance. Axial blade tip clearance may refer to a distance between the tip of the blade  22  and the turbine shroud  23  measured in the axial direction  30 . Radial blade tip clearance may refer to a distance between the tip of the blade  22  and the turbine shroud  23  measured along a radial direction  40 , generally perpendicular to the axial direction  30 . In the illustrated embodiment, an axial displacement control system  42  may be utilized for controlling axial displacement of rotor and/or stator components. In doing so, axial blade tip clearance may be reduced or negated, although this may, as set forth below, simultaneously reduce the radial blade tip clearance component as well. The axial displacement control system  42 , for example, may export a portion  44  of the compressed air  19  (or some other coolant, such as an inert gas (e.g., nitrogen, steam, vapor, water, refrigerant, etc.)) to a heat exchanger  46  (e.g., a direct heat exchanger and/or an indirect heat exchanger using a liquid or gas coolant), which may cool the portion  44  of compressed air  19 , generating cooled compressed air  48 . The cooled compressed air  48  may then be used to cool components of the rotor. For example, the cooled compressed air  48  may be used to cool the shaft  24  at locations within the transition shroud  38 . Alternatively or additionally, the cooled compressed air  48  may be used to cool compressor blades  28  proximate the control system  42 . Further, the cooled compressed air  48  may be used to cool the shaft  24  closer to the turbine  18  or may be used to cool rotor components proximate a connection of the blades  22  of the turbine  18  to the shaft  24 . However, in general, the cooled compressed air  48  may be directed to an area substantially defined upstream of the turbine  18 . Indeed, cooling components within the turbine  18  (e.g., the turbine blades  22  or discs thereof) or too far downstream within the turbine  18  may lead to the turbine blades  22  contracting radially away from the turbine shroud  23  toward the shaft  24 , increasing blade tip clearances. 
     By cooling components of the rotor, thermal expansion of the rotor components in the axial direction  30  may be reduced. Thus, the turbine blades  22  may be blocked from extending away from contact with the turbine shroud  23  (or honeycomb structure thereof) in the axial direction  30 . For example, by cooling the shaft  24 , thermal expansion of the shaft  24  in the axial direction  30  is reduced. Since the turbine blades  22  are coupled to the shaft  24 , the turbine blades  22  likewise are not displaced, or have a reduced displacement, in the axial direction  30 . Because the turbine shroud  23  gradually increases in cross-sectional area (e.g., a tapered annular wall) in the axial direction  30 , displacement of the turbine blades  22  in the axial direction  30  cause the turbine blades  22  to separate from the turbine shroud  23  (or honeycomb structure disposed on the turbine shroud  23 ). By blocking thermal expansion of the shaft  24 , separation of the turbine blades  22  from the honeycomb structure of the shroud  23  is reduced or eliminated. Further, because the cooled compressed air  48  is output from the heat exchanger  46  mostly to portions of the rotor of the turbomachine system  10 , as opposed to the stator, the stator (e.g., the turbine shroud  23  and the turbine vanes  25 ) may be allowed to thermally expand in the axial direction  30  toward the turbine blades  22 . Thus, the honeycomb structure disposed on the turbine shroud  23  or proximate the turbine shroud  23  may be axially displaced into tips of the turbine blades  22 . 
     It should be noted that the control system  42  may selectively utilize techniques described above based on certain operating conditions or stages of operation. For example, during certain operating intervals (e.g., stages of operation), it may be less beneficial to actively reduce or actively eliminate blade tip clearances than during other operation intervals. Indeed, during some operation intervals, blade tip clearances may be eliminated without the use of the control system  42  at all. Thus, the portion  44  of the compressed air  19  exported to the heat exchanger  46  may be exported to the heat exchanger  46  particularly during certain operating intervals (e.g., stages of operation) where blade tip clearances may benefit from rotor cooling. For example, the control system  42  may export the portion  44  of the compressed air  19  to the heat exchanger  46  for cooling rotor components when the turbomachine system  10  is at full speed no load, i.e., the turbomachine system  10  is running at full speed but is not coupled to the load  36 . Alternatively, the control system  42  may export the portion  44  of the compressed air  19  to the heat exchanger  46  for cooling rotor components during other intervals of operation, such as during all intervals of start-up between full speed no load and steady state operation. Further, depending on operating conditions (or stages of operation), the control system  42  may export a certain amount of compressed air  19  to the heat exchanger  46  and may cool the compressed air  19  to a certain extent depending on operational inputs taken into account by the control system  42 . The control system  42  and the various components which may be controlled via the control system  42  will be described in detail below, with reference to later figures. 
     Turning now to  FIGS. 2 and 3 , cross-sectional side views of one turbine blade  22  and a portion of the turbine shroud  23  is shown, taken within lines  2 - 2  of  FIG. 1 .  FIGS. 2 and 3  are intended to clarify certain aspects of blade tip clearances relative to components of the turbomachine system  10  proximate the blade tip clearance. Focusing on  FIG. 2 , a tip  70  of the blade  22  is shown slightly separated from a honeycomb structure  72  disposed on a portion of the turbine shroud  23 , where the honeycomb structure  72  is a softer material (e.g., adradable material) than the tips  70  of the blades  22 . For example, the honeycomb structure  72  may include any adradable material. The honeycomb structure  72  may include a base material having a nickel base foil (Nickel-16Chromium-4.5Aluminum-3.5Iron), with or without a gel aluminizing coating. Other embodiments of the honeycomb structure  72  may include a porous metallic material with polyester pore formers that are burned after ignition and mixed with metallic powders (e.g., MCrAlY or Cobalt/Nickel-Chromium-Aluminum-Yttrium), where the polyester pore formers may be applied via plasma spray. In some embodiments, soft metals such as Ni, Graphite, and/or Al may be used for the adradable material of the honeycomb structure  72 . Further, foam metals may be used. 
     In accordance with present embodiments, the honeycomb structure  72  may be conical or cylindrical in shape. For example, the illustrated honeycomb structure  72  is conical, such that axial thermal displacement of stator/rotor components may cause the blade tips  70  to move axially (e.g., opposite to direction  74 ) into the conical honeycomb structure  72 , or cause the turbine shroud  23  to move axially (e.g., in direction  74 ) into the blade tips  70 , as shown in  FIG. 3 . However, the honeycomb structure  72  may also be any other shape configured to enable the tips  70  of the blades  72  to cut into the honeycomb structure  72  during both transient and steady state operation. For example, some embodiments may include a cylindrical honeycomb structure  72  that is not sloped as shown in the illustrated embodiment. During transient operation, the blade tip  70  may carve out a trench in a certain portion of the honeycomb structure  72  (e.g., cylindrical honeycomb structure  72 ). During steady state operation, the blade tip  70  may be enabled to contact untrenched honeycomb (e.g., a different portion) of the honeycomb structure  72 , by way of stator and/or rotor axial thermal expansion control, in accordance with present embodiments. Accordingly, the honeycomb structure  72  may give way to the tips  70  of the blades  22  such that the blades  22  cut into the honeycomb structure  72 , during both transient and steady state operation. Thus, blade tip clearances are reduced during transient and steady state operation or loading. Further, the honeycomb structure  72  (e.g., the adradable material) generally enables rotation of the turbine blade  22  without exerting substantial resistance against the rotation of the turbine blade  22 . As previously described, the turbine blade  22 , during operation, may be rotating as a component of the rotor. In the illustrated embodiment, the turbine blade  22  may rotate in a first circumferential direction  74 , about the rotational axis  26 . 
     The illustrated tip  70  of the turbine blade  22  is separated from the honeycomb structure  72 , such that a clearance exists between the tip  70  and the honeycomb structure  72 . The clearance may include an axial component (e.g., an axial clearance  74 ) and a radial component (e.g., a radial clearance  76 ). The axial clearance  74  and the radial clearance  76  may both be eliminated or reduced in one of two ways. Moving the turbine blade  22  and the honeycomb structure  72  closer together in the axial direction  30 , such that the blade tip  70  and the honeycomb structure  72  come into contact, eliminates both axial clearance  74  and radial clearance  76 . Moving the turbine blade  22  and the honeycomb structure  72  closer together in the radial direction  40 , such that the blade tip  70  and the honeycomb structure  72  come into contact, also eliminates both axial clearance  74  and radial clearance  76 . Indeed, reduction of the blade tip clearances  74 ,  76  in both of the above described manners is made possible by the angled orientation of the honeycomb structure (e.g., tapered annular structure about axis  26 ) and the increasing cross-sectional area of the turbine shroud  23  in the axial direction  30 . 
     Embodiments of the present disclosure are concerned with utilizing the control system  42  to bring the honeycomb structure  72  and the turbine blade  22  tip  70  together in the axial direction  30 , although some thermal expansion and/or contraction of components may also occur in the radial direction  40 . This may be achieved by reducing or eliminating thermal expansion of the turbine blade  22  by cooling rotor components which the turbine blade  22  is coupled to, e.g., the shaft  24  (not shown) of the rotor. Alternatively or additionally, eliminating blade tip clearance may be achieved by effecting thermal expansion of the stator (e.g., the turbine shroud  23  of the stator) in the axial direction  30 , such that the honeycomb structure  72  disposed on the turbine shroud  23  may be axially displaced into the tip  70  of the turbine blade  22 . This may be achieved by the use of the control system  42 , as set forth in detail below, and may also be enhanced by selecting a low coefficient of thermal expansion material for the rotor (such that axial expansion of the rotor is reduced) and by selecting a high coefficient of thermal expansion material for the stator (such that axial expansion of the stator may be increased), at least relative to one another. The use of the control system  42  to achieve reduction or elimination of blade tip clearances, particularly through axial movement of components of the turbomachine system  10  in the axial direction  30 , will be described in detail below with reference to later figures. 
     Turning now to  FIG. 4 , a cross-sectional side view of a portion of an embodiment of the turbomachine system  10  is shown. The illustrated embodiment of the turbomachine system  10  includes the rotor comprising the shaft  24 , the compressor blades  28 , and the turbine blades  22 , along with a cooling area  80  (e.g., cooling channel or cooling cavity) running through a portion of the shaft  24  near a midsection  82  of the turbomachine system  10 . The cooling area  80  may be used to cool the shaft  24  (of the rotor) via the control system  42 , as previously described, and may be internal to the shaft  24 , external to the shaft  24 , or may include portions of both. Also included in the illustrated embodiment is the stator, comprising the compressor shroud  32 , the compressor vanes  34 , the turbine shroud  23 , and the turbine vanes  25 . The optional transition shroud  38  is also shown, although the transition shroud  38  may actually be a part of the compressor shroud  32  and/or a part of the turbine shroud  23 . Indeed, all three of the shrouds  23 ,  32 , and  38  may be one integral shroud used as a casing for the stator of the turbomachine system  10 . 
     As previously described, air (or other coolants, such as an inert gas (e.g., nitrogen, steam, liquid, vapor, etc.) external to the turbomachine system  10  is drawn into the compressor  12  and is compressed via the compressor vanes  34  and compressor blades  28  to generate compressed air  19 . The compressed air  19  is delivered to the combustors  14  (one shown), along with fuel from the fuel nozzle  16 . The combustor  14  combusts the compressed air  19  to generate combustion gases  20 , which are routed through the turbine blades  22  of the turbine  18  for driving the turbine blades  22  into rotation. The turbine blades  22  are coupled to the shaft  24 , such that the turbine blades  22  drive the shaft  24  into rotation, which, in turn, drives the compressor blades  28  into rotation. 
     Some of the compressed air  19  generated by the compressor  12  may be diverted from the combustors  14 . For example, the portion  44  of compressed air  19  is diverted away from the combustors  14  via the control system  42 . The control system  42  may include one or more sensors  84 , a controller  86 , and a valve  88 , where the one or more sensors  84  may be configured to detect pressure, temperature, light, vibration, noise, combustion dynamics, or a combination thereof, all of which may be configured to indicate a need to increase or decrease clearance. The controller  86  may be included with, or may be a part of, a processor and may include memory  90  with executable instructions stored on the memory  90 . For example, the controller  86  may include executable instructions which, when executed, determine if, when, and/or how much of the compressed air  19  may be diverted from the combustor  14 . The controller  86  may instruct the valve  88  to open fully, or to a certain degree, such that an appropriate amount of the compressed air  19  is diverted from the combustors  14 . Thus, the portion  44  of diverted compressed air  19  may be appropriately cooled via the heat exchanger(s)  46  and routed through or proximate rotor components (e.g., the shaft  24 ) for cooling the rotor components. 
     The controller  86  may accept input data from one or more of the sensors  84 , which may provide data to the controller  86  relating to operation conditions of the turbomachine system  10 . Operating conditions may include, for example, temperature of various components of the turbomachine system  10 , axial displacement measurements of various components of the turbomachine system  10  (e.g., the shaft  24 ), or stages of operation of the turbomachine system  10 . Stages of operation may include cold start (CS) (e.g., when the turbomachine system  10  is first started), full speed no load (FSNL) (e.g., when the turbomachine system  10  is at full speed but not connected to the load  36 ), full speed full load (FSFL) (e.g., when the turbomachine system  10  is at full speed and is just connected to the load  36 ), steady state (SS) (e.g., when the turbomachine system  10  is no longer in transient operation), shutdown, or some other transient or steady state stage or condition. The sensors  84  may also detect axial displacement of turbine components and provide data related to the axial displacement of the turbine components to the controller  86 . For example, one sensor  84  may be disposed on the shaft  24  proximate a third stage turbine blade  92  of the turbine  18 . The sensor  84  may detect axial displacement of the shaft  24  where the sensor  84  is located (e.g., proximate the third stage turbine blade  92 ) relative to a home position of the sensor  84  (e.g., in the axial direction  30 ). The home position of the sensor  84  (e.g., along the rotational axis  26 ) may be a location of the sensor  84  when the turbomachine system  10  is off-line. Thus, when the turbomachine system  10  begins to operate, the sensor  84  may detect axial displacement of the shaft  24  relative to the home position of the sensor  84  along the rotational axis  26  and relay information related to the axial displacement to the controller  86 . The controller  86  may then determine if, when, and/or how much of the compressed air  19  should be diverted to the cooling area  80  for cooling the shaft  24  or other rotor components proximate the cooling area  80 . Additionally, based on feedback from the sensors  84  (or based on some other input information), the controller  86  may determine when to block compressed air  19  from being diverted to the cooling area  80  for cooling the shaft  24  or other rotor components proximate the cooling area  80 , such as when the blade tips  70  are already contacting the honeycomb structure  72 . 
     The controller  86  is coupled to the sensors  84 , the valve  88  (e.g., via actuators or drivers), and/or the heat exchangers  46  (e.g., via valves or other controls). Thus, the controller  86  may control operation of any one or more of the sensors  84 , valve  88 , and heat exchangers  46 . The controller  86  may be capable of controlling any facet of the valve  88  (e.g., if and when to open the valve  88 , to what extent to open the valve  88 , etc.) and any facet of the heat exchangers  46  (e.g., to what extent to cool the diverted portion  44  of compressed air  19 ). The controller  86  may also be capable of receiving data input(s) from any one or more of the sensors  84  for determining how to appropriately control the valve  88  and/or heat exchangers  46 . The controller  86  may also receive a manual input from an operator. The controller  86  may be electrically coupled to the sensors  84 , valve  88 , and heat exchangers  46 , or the controller  86 , sensors  84 , valve  88 , and heat exchangers may be coupled to a network  96  (e.g., Internet, intranet, industrial control network, etc.) or other wired or wireless system, such that information and instructions may be shared between the components via the network  96 . Further, in some embodiments, the controller  86  and the valve  88  may be an integral component or physically coupled together in close proximity to one another. It should also be noted that, in some embodiments, the controller  86  may not be coupled to the heat exchanger(s)  46 . Accordingly, in some embodiments, the heat exchanger(s)  46  may cool the diverted portion  44  of compressed air  19  to the same extent at any time, once the portion  44  is allowed to pass through the valve  88 . 
     After determining that blade tip clearances may be reduced or eliminated via cooling of rotor components (in particular, the shaft  24 ), the controller  86  may open the valve  88 . The controller  86  may, for example, enable rotor cooling when the turbomachine system  10  is at a certain stage of operation. For example, after the turbomachine system  10  is at full speed no load (FSNL), blade tip clearances may be high or increasing, which enables hot combustion gases  20  to leak over the tips  70  of each turbine blade  22 . Accordingly, the controller  86  may enable rotor cooling after reaching FSNL by opening the valve  88 . The same may be true when the turbomachine system  10  is at full speed full load (FSFL), steady state (SS), or cold start (CS), or any other stage of operation, if conditions so permit. In general, the controller  86  is configured to control flow of coolant (e.g., compressed air, steam, refrigerant, or some other gas, liquid, or vapor) through the heat exchanger(s)  46 , which in turn controls a degree of cooling of components of the turbomachine  10 . 
     The portion  44  of compressed air  19  may then be diverted into the cooling area  80 , where the heat exchangers  46  cool the portion  44  of compressed air  19  to generate the cooled compressed air  48 . The cooled compressed air  48  may be routed through the cooling area  80 , which may be defined in part by one or more components of the rotor. In the illustrated embodiment, the cooling area  80  is defined entirely within the shaft  24  of the rotor, and extends below the combustor  14  from (or just beyond) a first end  97  of the combustor  14  to (or just beyond) a second end  98  of the combustor  14 . It should be noted that the first end  97  of the combustor  14  may be a first end of a chamber of the combustor  14 , but that other components of the combustor  14  (e.g., fuel injectors) may extend opposite the axial direction  30  beyond the cooling area  80 . Further, in the illustrated embodiment, the cooling area  80  includes portions directing the cooled compressed air  48  backwardly toward the compressor  12  (e.g., opposite the axial direction  30 ). The cooling area  80  also includes portions directing the cooled compressed air  48  forwardly toward the turbine  18  (e.g., in the axial direction  30 ). Further still, the cooling area  80  (e.g., cooling channel) may have outlets  99 , such that the cooled compressed air  48 , after extracting heat therefrom, for example, the shaft  24 , may exit the shaft  24  of the turbomachine  10 . In other embodiments, the cooling area  80  may be external to the shaft  24  and/or the cooling area  80  may contact or be disposed proximate other components of the rotor. Indeed, the cooling area  80  may be a channel, or a series of channels. Alternatively, the cooling area  80  may be an internal area of, for example, one or more rotor components, where the internal area may be defined by other features of the rotor component(s). 
     Further, the cooling area  80 , depending on the embodiment, may be disposed in particular locations of the turbomachine  10 . For example, in some embodiments, the cooling area  80  may be disposed proximate final stages (e.g., compressor blade  28  stages) of the compressor  12  and/or proximate initial stages (e.g., turbine blade  22  stages) of the turbine  18 . However, in some embodiments, the cooling area  80  may be disposed substantially proximate only rotor components, or mostly only rotor components, and in particular the shaft  24  of the rotor. Thus, the shaft  24  may be cooled, when appropriate, such that the shaft  24  is blocked from thermally expanding too much in the axial direction  30 . Otherwise, the blade tips  70  may be axially displaced in the axial direction  30 , away from the honeycomb structures  72 , such that blade tip clearances are increased. Further, the cooling area  80  may, in some embodiments, not extend very far into the turbine  18 , if at all, as cooling of rotor components within the turbine  18  (e.g., turbine blades  22 ) may radially contract the turbine blades  22  away from the turbine shroud  23  and toward the shaft  24 , which increases blade tip clearances. 
     It should be noted that it may be desirable, as described above, to enable cooling via the control system  42  at certain operating intervals or conditions in order to reduce blade tip clearances, but that it may also be desirable to block cooling of rotor components via the control system  42  to block a reduction in blade tip clearances in certain other operating intervals or conditions. Put differently, if the tips  70  are already cutting into the honeycomb structure  72 , it may be beneficial to block cooling such that the tips  70  do not eventually cut into a component radially outward from the honeycomb structure  72 , such as the turbine shroud  23 . For example, in one embodiment, during start up or shutdown (e.g., transient stages or conditions), it may be beneficial to block coolant from cooling rotor components. During steady state stages or conditions, it may be beneficial to enable coolant to cool rotor components. Alternatively, in another embodiment, during start up or shutdown (e.g., transient stages or conditions), it may be beneficial to enable coolant to cool rotor components. In such an embodiment, during steady state stages or conditions, it may be beneficial to block coolant from cooling rotor components. 
     In some embodiments, the cooling area  80  may be disposed proximate some stator components. However, in general, the cooling area  80  is disposed mostly proximate rotor components. Indeed, blade tip clearances may be further reduced by ensuring that the stator, as described above, and in particular the turbine shroud  23  and the honeycomb structures  72  disposed on the turbine shroud  23 , thermally expands in the axial direction  30 , into or toward the tips  70  of the turbine blades  22 . Indeed, as indicated by line  100  in the illustrated embodiment, the turbine shroud  23  (and the turbine  18  in general) opens up in the axial direction  30  along the rotational axis  26 . In other words, the line  100  (e.g., slope) extending through the turbine shroud  23  is sloped relative to the rotational axis  26 , such that a cross-sectional area of the turbine  18  increases in the axial direction  30 . Thus, blade tip clearances may be reduced or eliminate by axially displacing, or preventing axial displacement, of certain components due to the slope  100  of the turbine shroud  23 . For example, by contracting or blocking axial displacement (e.g., by cooling) of the shaft  24  in the axial direction  30  and, thus, the blades  22  (having blade tips  70 ) coupled to the shaft  24 , the blade tips  70  are blocked from separating from the honeycomb structures  72  disposed along the slope  100  of the turbine shroud  23 . Further, by effecting axial displacement (e.g., through thermal expansion) of the turbine shroud  23  in the axial direction  30 , the turbine shroud  23 , since it is sloped along line  100 , thermally expands into or toward the blade tips  70 . By controlling rotor cooling and stator heating (e.g., via the control system  42 ) either simultaneously or independently during various stages of operation, blade tip clearances may be reduced or eliminated, when appropriate. 
     It should be noted, however, that the honeycomb structure  72  may or may not follow the slope  100 . For example, in the illustrated embodiment, the honeycomb structure  72  is conical in accordance with the description above. However, in some embodiments, the honeycomb structure  72  may be cylindrical. In such embodiments, the blade tips  70  may contact a first portion of the honeycomb structure  72  during transient loading, and a second, untrenched portion of the honeycomb structure during steady state loading. The blade tips  70  may contact different portions of the honeycomb structure  72  via axial thermal displacement (e.g., via cooling/heating) of stator and/or rotor components, in accordance with the present disclosure. 
     Turning now to  FIG. 5 , a process flow diagram of a method  110  for reducing blade tip clearances is shown. The illustrated method  110  includes generating compressed air  19  (block  112 ) and diverting the portion  44  of the compressed air  19  to the heat exchanger  46  (block  114 ). The compressed air  19  may be generated by the compressor  12  of the turbomachine system  10  and the portion  44  of compressed air  19  may be diverted to the heat exchanger  46  via the valve  88 , as previously described, which may be controlled by the controller  86 . The method  110  further includes cooling the portion  44  of the compressed air  19  via the heat exchanger  46  to generate cooled compressed air  48  (block  116 ). Further still, the method  110  includes routing the cooled compressed air  48  through an area of the turbomachine system  10  for cooling rotor components of the turbomachine system  10  (block  118 ). The area is disposed proximate the rotor components and extends proximate the compressor  12  of the turbomachine system  10 . The area is disposed proximate the rotor components such that the rotor components may be cooled, which reduces an axial displacement of the rotor components. Reducing the axial displacement of the rotor components may reduce blade tip clearances between turbine blades  22  and the turbine shroud  23  (or the honeycomb structure  72  disposed on the turbine shroud  23 ). 
     In accordance with the present disclosure, decreasing blade tip clearances via controlling axial displacement of components of the turbomachine system  10  may reduce leakage of combustion gases over the tips  70  of the turbine blades  22 . Further, utilizing the presently discloses control system  24  to do so, as opposed to using a hydraulic or actuation displacement mechanism, may save material cost and complexity of manufacturing. Further still, by ensuring that cooling of rotor components does not extend too far into the turbine  18 , the rotor components may be blocked from thermal expansion in the axial direction  30  while the turbine blades  22  do not contract away from the turbine shroud  23  toward the shaft  24 . 
     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 have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.