Patent Publication Number: US-7708518-B2

Title: Turbine blade tip clearance control

Description:
FIELD OF THE INVENTION 
     The invention relates in general to turbine engines and, more particularly, to blade tip clearances in the turbine section of a turbine engine. 
     BACKGROUND OF THE INVENTION 
       FIG. 1  shows a cross-section through a portion of a turbine engine. A turbine engine  10  can generally include a compressor section  12 , a combustor section  14  and a turbine section  16 . A centrally disposed rotor  18  can extend through the three sections. 
     Generally, the combustor section  14  is enclosed within a casing  20  that can form a chamber  22 , together with the aft end of the compressor casing  24  and a housing  26  that surrounds a portion of the rotor  18 . A plurality of combustors  28  and ducts  30  can be provided within the chamber  22 , such as in an annular array about the rotor  18 . Each duct  30  can connect one of the combustors  28  to the turbine section  16 . 
     The turbine section  16  can include an outer casing  32  which encloses alternating rows of stationary airfoils  34  (commonly referred to as vanes) and rotating airfoils  36  (commonly referred to as blades). Each row of blades can include a plurality of airfoils  36  attached to a disc  38  provided on the rotor  18 . The rotor  18  can include a plurality of axially-spaced discs  38 . The blades  36  can extend radially outward from the discs  38  and terminate in a region known as the blade tip  40 . 
     Each row of vanes can be formed by attaching a plurality of airfoils  34  to the stationary support structure in the turbine section  16 . For instance, the airfoils  34  can be hosted by a vane carrier  42  that is attached to the outer casing  32 . The vanes  34  can extend radially inward from the vane carrier  42  or other stationary support structure to which they are attached. 
     In operation, the compressor section  12  can induct ambient air and can compress it. The compressed air  44  from the compressor section  12  can enter the chamber  22  and can then be distributed to each of the combustors  28 . In the combustors  28 , the compressed air can be mixed with the fuel introduced through a fuel nozzle  46 . The air-fuel mixture can be burned, thereby forming a hot working gas  48 . The hot gas  48  can flow through the ducts  30  and then through the rows of stationary airfoils  34  and rotating airfoils  36  in the turbine section  16 , where the gas  48  can expand and generate power that can drive the rotor  18 . The expanded gas  50  can then be exhausted from the turbine  16 . 
     It should be noted that each row of blades  36  is surrounded by the stationary support structure of the turbine, which can be the outer casing  32 , the vane carrier  42  or a ring seal (not shown). The space between the blade tips  40  and the neighboring stationary structure is referred to as the blade tip clearance C. During engine operation, gas leakage can occur through the blade tip clearances C, resulting in measurable engine performance decreases in power and efficiency. 
     While small blade tip clearances C are desired to minimize gas leakage, it is critical to maintain a clearance C between the rotating turbine components (blades  36 , rotor  18 , and discs  38 ) and the stationary turbine components (vanes  34 , outer casing  32 , vane carriers  42  and ring seals) at all times. Rubbing of any of the rotating and stationary components can lead to substantial component damage, performance degradation, and extended outages. 
     However, during transient conditions such as during engine startup or part load operation, it can be difficult to ensure that adequate blade tip clearances C are maintained because the rotating parts and the stationary parts thermally expand at different rates. For instance, in a cold start situation, the rate of thermal expansion of the thermal stationary support structure is at least initially less than the rate of thermal expansion of the rotating turbine components due to the relatively larger size and thickness of the stationary support structure. As a result, the blade tip clearances C can actually decrease because the rotating components expand radially outward faster than the stationary support structure, raising concerns of blade tip rubbing. 
     To avoid blade tip rubbing, large tip clearances are initially provided so that minimum blade tip clearances C are maintained at known pinch points, that is, during operational conditions where the clearances C would otherwise be expected to be the smallest (hot restart, spin cool, etc.). However, because the minimum blade tip clearances C are sized for these pinch point conditions, the clearances C eventually become overly large as the rate of thermal expansion of the rotating components slows or substantially stops while the stationary support structure continues to grow radially outward. Such oversized clearances C can occur as the engine approaches or attains steady state operation, such as at base load. Consequently, engine power and efficiency can be reduced. 
     Thus, there is a need for a system that can improve engine performance by minimizing turbine tip clearances at desired engine operating conditions. 
     SUMMARY OF THE INVENTION 
     In one respect, aspects of the invention are directed to a method for controlling blade tip clearances in a turbine engine. The turbine engine has a compressor section, a combustor section, and a turbine section. The combustor section receives compressed air from the compressor section. The turbine section includes a rotor with a plurality of discs thereon. A plurality of blades are attached to each disc. Each blade extends radially outward from the disc to a blade tip. The blade tips are substantially proximate a stationary support structure surrounding the blades. The stationary support structure can be a vane carrier, a ring seal and/or an outer casing. The stationary support structure is at a first temperature. A blade tip clearance is defined between the blade tips and the stationary support structure. 
     According to the method, a portion of the compressed air from the combustor section is extracted. Next, the extracted portion of air is cooled to a second temperature that is less than the first temperature. At least a portion of the cooled air at the second temperature is then passed in heat exchanging relation with the stationary support structure such that the stationary support structure thermally contracts. Such contraction can cause the blade tip clearance to decrease. The passing step can be selectively performed upon the occurrence of an operational parameter. The method can also involve measuring the blade tip clearance and selectively performing the passing step to ensure a target blade tip clearance is maintained. 
     In one embodiment, at least the passing step can be performed during substantially steady state engine operation. In one embodiment, at least the passing step can be performed during base load operation. In yet another embodiment, at least the passing step can be performed during part load operation. 
     The method can also include the step of routing the air that has passed in heat exchanging relation with the stationary support structure back to the air at the second temperature so as to form an air mixture at a mixture temperature. The mixture temperature can be measured and, when the measured mixture temperature exceeds a predetermined temperature, the cooling step can be adjusted such that the extracted portion of air is cooled to a temperature less than the second temperature. 
     In another respect, aspects of the invention related to a blade tip clearance control system. The system includes a turbine engine having a compressor section, a combustor section having a chamber receiving compressed air from the compressor section, and a turbine section. The turbine section includes a plurality of discs mounted to a rotor. A plurality of blades are attached to the discs; each blade extends radially outward from the disc to a tip. The system also includes stationary support structure substantially surrounding at least a portion of the blades. A clearance is defined between the tips of the blades and the stationary support structure. The stationary support structure is at a first temperature. The stationary support structure can be one or more of the following: a vane carrier, a ring seal and an outer casing. 
     The system further includes a rotor cooling air circuit that includes a fluid conduit and a cooler disposed along the fluid conduit. The fluid conduit is connected in fluid communication with the chamber of the combustor section such that a portion of the compressed air in the chamber is received within the fluid conduit. The portion of compressed air passes in heat exchanging relation with the cooler such that the temperature of the portion of air is reduced to a second temperature that is less than the first temperature. 
     A supply conduit is connected in fluid communication with the fluid conduit and extends therefrom. The supply conduit routes at least a portion of the air at the second temperature to the stationary support structure so that the air passes in heat exchanging relation with the stationary support structure. As a result, the stationary support structure contracts to reduce the clearance. In one embodiment, one or more passages extend through at least a portion of the stationary support structure. The passage has an inlet and an outlet. The supply conduit is connected in fluid communication with the inlet of the passage such that the passage receives the air at the second temperature. 
     In one embodiment, the system can also include a return conduit positioned to receive the air that has passed in heat exchanging relation with the stationary support structure. The return conduit can be connected in fluid communication with the fluid conduit, downstream of the area where the supply conduit connects to the fluid conduit. Thus, air that has passed in heat exchanging relation with the stationary support structure can be routed back to the fluid conduit. A temperature measurement device can be operatively associated with the fluid conduit downstream of the area where the return conduit connects to the fluid conduit. 
     A valve can be operatively positioned along one of the fluid conduit and the supply conduit to selectively permit and prohibit the supply of air at the second temperature to the stationary support structure. In one embodiment, the valve can permit the air at the second temperature to be supplied to the stationary support structure during base load engine operation. 
     In yet another respect, aspects of the invention concern a blade tip clearance control system. The system includes a turbine engine that has a compressor section, a combustor section having a chamber receiving compressed air from the compressor section, and a turbine section. The turbine section including a plurality of discs mounted to a rotor. A plurality of blades are attached to the discs, and each blade extends radially outward therefrom to a tip. 
     A stationary support structure substantially surrounds at least a portion of the blades. The stationary support structure can be a vane carrier, a ring seal, an outer casing or any combination thereof. A clearance is defined between the tips of the blades and the stationary support structure. The stationary support structure has one or more passages extending therethrough. The passage has an inlet end and an outlet end. The stationary support structure is at a first temperature. 
     The system includes a rotor cooling air circuit with a fluid conduit and a cooler disposed along the fluid conduit. The fluid conduit connects between and in fluid communication with the chamber of the combustor section and the inlet end of the passage. A portion of the compressed air in the chamber is received within the fluid conduit and passes in heat exchanging relation with the cooler such that the temperature of the portion of air is reduced to a second temperature, which is less than the first temperature. 
     A supply conduit connects between and in fluid communication with the second conduit and the inlet end of the passage. The supply conduit routes at least a portion of the air at the second temperature to the passage; the air passes through the passage in heat exchanging relation with the stationary support structure. Thus, the stationary support structure contracts to reduce the clearance. A valve is operatively positioned along one of the fluid conduit and the supply conduit to selectively permit and prohibit the supply of air at the second temperature to the stationary support structure. 
     In one embodiment, a return conduit can connect between and in fluid communication with the outlet end of the passage and the fluid conduit. The return conduit can connect to the fluid conduit downstream of the area where the supply conduit connects to the fluid conduit. Thus, air exiting the passage is routed back to the rotor cooling air circuit. A temperature measurement device can be operatively associated with the fluid conduit downstream of the area where the return conduit connects to the fluid conduit. The temperature measurement device can be operatively connected to the cooler, allowing the temperature of the coolant exiting the cooler can be altered as necessary. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a cross-sectional view through a portion of a known turbine engine. 
         FIG. 2  is a partial cross-sectional view of a blade tip clearance control system according to aspects of the invention, several engine components not shown for purposes of clarity. 
     
    
    
     DETAILED DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     Aspects of the present invention relate to a system and method for controlling blade tip clearances in the turbine section of the engine. Embodiments of the invention will be explained in the context of one clearance control system, but the detailed description is intended only as exemplary. Embodiments of the invention are shown in  FIG. 2 , but aspects of the invention are not limited to the illustrated structure or application. 
     Generally, the clearance control system according to aspects of the invention involves passing a fluid in heat exchanging relation with the vane carrier  42  or other stationary support structure that is proximate the tips  40  of the rotating airfoils  36 . Because air is readily available in a turbine engine, aspects of the invention are particularly suited for using air as the fluid. More specifically, the blade tip clearance control system according to aspects of the invention can make use of the compressed air  44  from the chamber  22  in the combustor section  14 . 
     As is known, the compressed air  44  from the compressor  12  can be used to cool the rotor  18  or to internally cool the turbine blades  36 , among other things. Referring to  FIG. 1 , a portion  52  of the compressed air  44  from the compressor  12  can be extracted from the chamber  22  and routed externally of the engine  10  through a fluid conduit  53  connected in fluid communication with the chamber  22 . The fluid conduit  53  can be a single conduit or a plurality of conduit segments. For convenience, the fluid conduit  53  will be described herein as including a first conduit segment  54  and a second conduit segment  60 , but it will be understood that aspects of the invention are not limited to such an arrangement. 
     By entering the first conduit segment  54 , the portion of air  52  bypasses the combustors  28 . The portion of air  52  can be cooled by an external cooler  56  disposed along the fluid conduit  53 . In one embodiment, the cooler  56  can be a fin-fan heat exchanger. Alternatively, the cooler  56  can be a kettle boiler, which can be used to generate steam in the bottoming cycle in a combined cycle power plant. However, aspects of the invention are not limited to any particular cooler  56 , which can be almost any type of heat exchanger. The cooler  56  can be used to reduce the temperature of the portion of air  52 . In one embodiment, the temperature of the air  52  extracted from the chamber  22  can be about 800 degrees Fahrenheit. In such case, the cooler  56  can be used to reduce the temperature of the air  52  to about 400 degrees Fahrenheit. These temperatures are provided as examples, and it will be understood that these temperatures can vary from system to system. 
     After exiting the cooler  56 , the cooled air  58  can flow along the fluid conduit  53 , such as the second conduit segment  60 . The fluid conduit  53  can route the cooled air to one or more openings  62  formed in the housing  26 , thereby allowing the air  58  to enter a cooling air manifold  64  that surrounds a portion of the rotor  18 . From there, the cooling air  58  can be used to cool various engine components. For convenience, the above-described system of cooling extracted air  52  from the chamber  22  and redirecting it toward the rotor  18  will be generally referred to herein as the rotor cooling air circuit  66 . 
     According to aspects of the invention, blade tip clearances C can be affected by passing at least a portion of the cooled air  58  in the rotor cooling air circuit  66  in heat exchanging relation with the vane carrier  42  or other stationary support structure surrounding one or more rows of blades  36  in the turbine section  16 . Greater control of the blade tip clearance C can be achieved by selectively passing the cooled air  58  in heat exchanging relation with the vane carrier  42  or other stationary support structure surrounding one or more rows of blades  36  in the turbine section  16 . One example of a blade tip clearance control system according to aspects of the invention is shown in  FIG. 2 . 
     While the cooled air can  58  can exchange heat with one or more of the components forming the stationary support structure. The following discussion will concern the vane carrier  70 , though it will be understood that aspects of the invention are not limited to the vane carrier  70 . It should be noted that the vane carrier  70  can be generally cylindrical in conformation. The vane carrier  70  can be a single piece, or the vane carrier  70  can be a plurality of substantially circumferentially adjacent segments. The term circumferentially is intended to mean circumferential relative to the turbine. In one embodiment, the vane carrier  70  can be made of two generally semi-cylindrical portions. It will be understood that aspects of the invention can be applied to any vane carrier  70  regardless of the configuration and that the term “vane carrier” as used herein refers to any of such configurations. 
     The vane carrier  70  can be configured to exchange heat with the cooled air  58  from the rotor cooling air circuit  66 . In one embodiment, the cooled air  58  can be passed in heat exchanging relation with at least a portion of the exterior of the vane carrier  70 . Thus, the exterior of the vane carrier  70  can be configured as needed to facilitate the exchange of heat between the vane carrier  70  and the cooled air  58 . 
     In another embodiment, the cooled air  58  can be passed in heat exchanging relation with at least a portion of the interior of the vane carrier  70 . Thus, the vane carrier  70  or other stationary support structure can be configured to receive a portion of air from the rotor cooling air circuit  66 . For example, at least one passage  72  can extend through the vane carrier  70  for receiving at least a portion of air  58  from the rotor cooling air circuit  66  and allowing it to flow through the vane carrier  70 . The passage  72  can extend between an inlet end  74  and an outlet end  76 . Preferably, a substantial portion of the passage  72  can extend generally in the axial direction relative to the turbine. Ideally, the passage  72  spans a substantial portion of the axial length of the vane carrier  70 . The passage  72  can be provided in the vane carrier  70  by, for example, machining or casting. 
     There can be any number of passages  72  in the vane carrier  70 , and embodiments of the invention are not limited to any particular number of passages  72 . For example, there can be a single passage  72  extending through the vane carrier  70 . In another embodiment, there can be two or more passages  72  in the vane carrier  70 . In cases where multiple passages  72  are provided, the passages  72  can be substantially equally or unequally circumferentially spaced about the vane carrier  70 . Further, the passages  72  can be substantially parallel to each other or at least one of the passages  72  can be non-parallel to the other passages  72 . Each passage  72  can be substantially straight or at least one passage  72  can be curved, bent, serpentine or otherwise non-straight. 
     The passage  72  can have any of a number of cross-sectional shapes. In one embodiment, the passage  72  can be substantially circular. However, the passage  72  can also be oval, rectangular, and polygonal, just to name a few possibilities. The cross-section area of the passage  72  can be substantially constant, or it can vary along the length of the passage  72 . In the case of multiple passages  72 , the passages  72  can have substantially identical cross-sectional geometries and areas, but at least one of the passages  72  can be different in any of the above respects. Each passage  72  can be sized as needed. 
     According to aspects of the invention, at least a portion of air  78  can be routed from the rotor cooling air circuit  66  and delivered to the vane carrier  70 . A supply conduit  80  can be connected in fluid communication with the second conduit segment  60  and, for example, the inlet end  74  of the passage  72  in the vane carrier  70 . The supply conduit  80  can be connected to the vane carrier  70  and the rotor cooling air circuit  66  in various ways, such as by fasteners, couplings, seals, adhesives and/or threaded engagement. The supply conduit  80  can be connected to the rotor cooling air circuit  66  almost anywhere along the second conduit segment  60 . Preferably, the supply conduit  80  connects to a portion of the second conduit segment  60  that is inside of the chamber  22 . The supply conduit  80  can be routed as needed within the chamber  22  to avoid interferences with other components and to minimize disruptions in the flow of air within the chamber  22 . 
     In some instances, a return conduit, such as a return conduit  82 , can be extend between and can be connected in fluid communication with the second conduit segment  60  and the outlet end  76  of the passage  72  in the vane carrier  70 . The return conduit  82  preferably connects to the second conduit segment  60  downstream (relative to the direction of the airflow in the second conduit segment  60 ) of where the supply conduit  80  connects to the second conduit segment  60 . The return conduit  82  can be connected to the second conduit segment  60  and the outlet end  76  of the passage  72  in the vane carrier  70  in various ways, such as by fasteners, couplings, seals, adhesives and/or threaded engagement. The return conduit  82  can be routed as necessary to avoid interferences with other components and to minimize disruptions in the flow of air within the chamber  22 . 
     The supply and return conduits  80 ,  82  can be sized as needed. The pipes  80 ,  82  can have any cross-sectional area such as circular, rectangular, triangular or polygonal. The cross-sectional area of each of the pipes  80 ,  82  can be substantially constant or it can vary. The pipes  80 ,  82  can be substantially straight, or they can include any number of bends, turns, curves, etc. The supply and return conduits can be defined by a single pipes  80 ,  82 , or they can be defined by a plurality of pipe segments (not shown). 
     While  FIG. 2  shows the inlet end  74  of the passage  72  located near the axial downstream end  84  of the vane carrier  70  and the outlet end  76  of the passage  72  located near the axial upstream end  86  of the vane carrier  70 , it will be understood that aspects of the invention are not limited to this arrangement. For example, it will be readily appreciated that the opposite arrangement can be provided, that is, the inlet end  74  of the passage  72  can be provided near the axial upstream end  86  of the vane carrier  70 , and the outlet end  76  of the passage  72  can be provided near the axial downstream end  84  of the vane carrier  70 . 
     Just as there can be any number of passages  72  in the vane carrier  70 , there can any number of supply and return conduits  80 ,  82 . It should be noted that the number of supply conduits  80  may or may not be equal to the number of return conduits  82 . Moreover, any number of supply and return conduits  80 ,  82  can be in fluid communication with any number of passages  72  in the vane carrier  70 . In one embodiment, each passage  72  in the vane carrier  70  can have a dedicated supply conduit  80  and/or a dedicated return conduit  82 . Alternatively, one supply conduit  80  can be in fluid communication with more than one passage  72  in the vane carrier  70 . For instance, the supply conduit  80  can include a plurality of branches (not shown) with each branch in fluid communication with the inlet end  74  of a respective passage  72 . In another embodiment, the supply conduit  80  can be in fluid communication with a plurality of passages  72  by way of a supply plenum (not shown) in the vane carrier  70 . The supply plenum can be in fluid communication with a plurality of passages  72 . 
     Similar arrangements can be provided between the return conduit  82  and the outlet end  76  of the passage  72 . For instance, the vane carrier  70  can include a return plenum (not shown) that allows fluid communication between the return conduit  82  and a plurality of passages  72 . There can be any number of supply and/or return plenums. The plenums can extend substantially circumferentially through at least a portion of the vane carrier  70 . The plenums can have various cross-sectional geometries and surface contours, such as those discussed above in the context of the passages  72 . 
     As noted earlier, greater control of the blade tip clearances can be achieved by selectively supplying and restricting air  78  to the vane carrier  70  or other stationary support structure. To that end, a system according to aspects of the invention can further include a flow regulator, such as a valve  88 . The valve  88  can be disposed anywhere along the supply conduit  80  and/or the second conduit segment  60  of the rotor cooling air circuit  66 . The valve  88  can be used to selectively permit and prohibit the flow of the air  78  from the rotor cooling air circuit  66  to the passage  72  in the vane carrier  70 . The valve  88  can be operated manually or by a controller (not shown) operatively associated with the valve  88 . The valve  88  can be any suitable valve. 
     Having described several of the individual components of a system according to aspects of the invention, one manner of using the blade tip clearance control system  68  will now be described. The method described herein is merely an example. Not every step described need occur, and the steps described are not limited to performance in the sequence described. For purposes of this example, it will be assumed that the operation begins from a cold start condition. 
     The turbine engine  10  can be operated as is known. From startup, the valve  88  can be closed so as to substantially restrict the air  58  in the second conduit segment  60  from entering the passage  72  in the vane carrier  70  and/or the supply conduit  80 . The valve  88  can remain closed until a desired first operational parameter is reached. The operational parameter can be, for example, substantially steady state operation including base load operation. The first operational parameter can be any condition where most of the components that can affect the blade tip clearance C (blades  36 , rotor  18 , discs  38 , outer casing  32 , vane carrier  70 , etc.) have thermally grown to their final shapes. In any of the above examples, the occurrence of the first operational parameter can be determined in various ways, such as by measuring engine power output. In one embodiment, the first operational parameter can occur when the engine is operating at about 90 percent power or greater. Alternatively, the first operational parameter can be a certain blade tip clearance C. To that end, blade tip clearances C can be measured during engine operation using sensors or probes, as is known. In still another embodiment, the first operational parameter can be the temperature of the stationary support structure, as measured by a thermal sensor or other temperature measurement device. 
     Once the desired operational parameter is reached, the valve  88  can be opened to allow at least a portion of the air  78  in the rotor cooling circuit  66  to be diverted therefrom. The air  78  can be directed to the vane carrier  70  by the supply conduit  80 . The temperature of the air  78  supplied to the vane carrier  70  will be less than the operational temperature of the vane carrier  70 . The air  78  can enter and travel through the passage  72  in heat exchanging relation with the vane carrier  70 . Consequently, the temperature of the vane carrier  70  will decrease, and the temperature of the air  78  will increase. The vane carrier  70  will thermally contract at least in the radial direction. This contraction causes the vane carrier  70  to move closer to the blade tips  40 , thereby reducing the blade tip clearance C. Thus, fluid leakage through the clearance C can be minimized and engine power and efficiency can be increased. 
     After passing in heat exchanging relation with the vane carrier  70 , the air  90  can be directed to various areas. In one embodiment, the air  90  can be routed back to the rotor cooling air circuit  66  by the return conduit  82 . The returned air  90  can mix with the cooled air  58  in the second conduit segment  60 . It will be appreciated that the temperature of the returning air  90  will be greater than the temperature of the air  58  in the second conduit segment  60 . As a result, the temperature of the air mixture in the rotor cooling air circuit  66  can be greater than the temperature of the air exiting the cooler  60 , which can have an impact on the intended downstream cooling uses. 
     Such temperature changes can be monitored with a temperature measurement device operatively positioned along the pipe  66  downstream of the point at which the return conduit  82  connects to the second conduit segment  56 . The temperature measurement device can be, for example, a thermocouple  92 . The temperature measurement device can be operatively connected to the cooler by way of a controller (not shown), which can alert an operator when the temperature of the rotor cooling air increases beyond a predetermined temperature. The controller can be, for example, a computer. Any undesired increases in the temperature of the rotor cooling air  58  can be corrected by changing the operating parameters of the cooler  56  so as to lower the temperature of the air exiting the cooler  56 . 
     While air  78  is being passed in heat exchanging relation with the vane carrier  70 , the blade tip clearance C can be monitored to prevent blade tip rubbing from occurring. The blade tip clearance C can be measured in any of the various manners known in the art, such as probe measurement. The blade tip clearance C can be actively adjusted, as needed, by selectively increasing and decreasing the amount of air  78  delivered to the vane carrier  70 , such as by way of the valve  88 . Thus, a target blade tip clearance can be maintained. The target blade tip clearance can be, for example, a minimum clearance, a preferred clearance or range of clearance. 
     The supply of air  78  to the vane carrier  70  can continue for so long as needed or is desired or when a second operational parameter is reached. In such case, air flow to the supply conduit  80  and/or to the passage  72  can be substantially restricted, such as by closing the valve  88 . The second operational parameter can be, for example, a minimum design blade tip clearance C. Alternatively, the second operational parameter can be part load operation, as measured by engine power output. In one embodiment, the second operational parameter can occur when the engine is operating at less than about 90 percent power. The second operational parameter can also be the temperature of the stationary support structure, as measured by a thermal sensor or other temperature measurement device. 
     While especially suited for minimizing the tip clearance in the first row of turbine blades in a turbine section of a turbine engine, aspects of the invention can be applied to any and all rows of blades in the turbine section. Further, as noted above, aspects of the invention can be particularly beneficial during steady state engine operation, such as at base load. However, aspects of the invention can be used during part load operation as well or any condition in which improved engine performance is desired. Thus, it will of course be understood that the invention is not limited to the specific details described herein, which are given by way of example only, and that various modifications and alterations are possible within the scope of the invention as defined in the following claims.