Abstract:
A method of operating a gas turbine engine is provided. The gas turbine engine includes at least one engine casing and at least one rotor assembly. The method includes directing airflow through a supply pipe and into a heat exchanger, lowering the temperature of the airflow in the heat exchanger, and directing the cooled airflow into the engine casing to cool the casing.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to gas turbine engines, and more particularly, to methods and apparatus to control gas turbine engine rotor assembly tip clearances during transient operation.  
         [0002]     Gas turbine engines typically include an engine casing that extends circumferentially around a compressor, and a turbine including a rotor assembly and a stator assembly. The rotor assembly includes at least one row of rotating blades that extend radially outward from a blade root to a blade tip. A radial tip clearance is defined between the rotating blade tips and a shroud attached to the engine casing.  
         [0003]     During engine operation, the thermal environment in the engine varies and may cause thermal expansion or contraction of the rotor and stator assemblies. This thermal growth or contraction may or may not occour uniformly in magnitude or rate. As a result, inadvertent rubbing between the rotor blade tips and the engine casing may occur or the radial clearances may be more open than the design intent. Continued rubbing between the rotor blade tips and engine casing may lead to premature failure of the rotor blade or larger clearances at other operating conditions which can result in loss of engine performance.  
         [0004]     To facilitate optimizing engine performance and to minimize inadvertent rubbing between the rotor blade tips and the engine casing, at least some known engines include a clearance control system. The clearance control system supplies cooling air to the engine casing to control thermal growth of the engine casing to facilitate minimizing inadvertent blade tip rubbing. The case is heated and cooled by air coming from fan, booster, or compressor compressor bleed sources, which may cause the case to shrink or expand due to changes in temperature.  
         [0005]     Some known clearance control systems account for disk elastic deflection and blade thermal growth from idle conditions to aircraft take-off by having a large clearance at idle in order to prevent blade tip rubs later in the engine cycle. These systems require a large change in temperature at steady-state conditions to reduce clearance to a minimum level. Typically, the temperature change of the case that is necessary to reduce the steady-state clearance is beyond the capability of these systems.  
       BRIEF DESCRIPTION OF THE INVENTION  
       [0006]     In one aspect, a method of operating a gas turbine engine is provided. The gas turbine engine includes an engine casing and at least one rotor assembly. The method includes directing airflow through a supply pipe and into a heat exchanger, lowering the temperature of the airflow in the heat exchanger, and directing the cooled airflow into the engine casing to cool the casing.  
         [0007]     In another aspect, a clearance control system for a gas turbine engine is provided. The gas turbine engine includes a compressor having at least one stage and a discharge, a high pressure turbine having at least one disk, and at least one engine casing extending circumferentially around the compressor and high pressure turbine. The clearance control system includes an air supply pipe configured to direct air from at least one of the fan booster and compressor to the high pressure turbine, and a heat exchanger in flow communication with the air supply pipe to cool an airflow passing through the air supply line.  
         [0008]     In another aspect, a gas turbine engine is provided that includes a compressor, a high pressure turbine, at least one engine casing, including a cooling air inlet, extending circumferentially around the compressor and high pressure turbine, and a clearance control system. The clearance control system includes an air supply pipe coupled to the casing cooling air inlet, and a heat exchanger coupled to and in flow communication with the air supply pipe to cool an airflow in the air supply pipe.  
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0009]      FIG. 1  is a schematic illustration of a gas turbine engine.  
         [0010]      FIG. 2  is an enlarged sectional schematic illustration of a portion of the gas turbine engine shown in  FIG. 1 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0011]     A clearance control system for a gas turbine engine that includes a heat exchanger to lower the temperature of air that is used to cool the turbine casing is described below in detail. The cooling air can come from any source inside the engine, for example, from the middle stages of the compressor, or the compressor discharge. The cooling air that has been cooled in the heat exchanger is used to cool the turbine casing and turbine shrouds. The air can also be directed to the high pressure turbine disk cavity to cool the aft surface of the high pressure turbine disk. Further, to improve efficiency, a portion of the cooling air is redirected to the heat exchanger to be cooled and reused. Using air that has been cooled by the heat exchanger permits the stator to achieve a greater change in temperature for clearance closure, especially during steady state conditions. Also, cooling the high pressure turbine disk reduces disk thermal growth, which typically accounts for the majority of the total closure of blade tip clearances. The clearance control system described in detail below permits tighter build clearances, reduced operational thermal closure of clearances, and minimizes blade tip rubs.  
         [0012]     Referring to the drawings,  FIG. 1  is a schematic illustration of a gas turbine engine  10  that includes, in an exemplary embodiment, a fan assembly  12  and a core engine  13  including a high pressure compressor  14 , and a combustor  16 . Engine  10  also includes a high pressure turbine  18 , a low pressure turbine  20 , and a booster  22 . Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disk  26 . Engine  10  has an intake side  28  and an exhaust side  30 . In one embodiment, the gas turbine engine is a GE90 available from General Electric Company, Cincinnati, Ohio. Fan assembly  12  and low pressure turbine  20  are coupled by a first rotor shaft  31 , and compressor  14  and high pressure turbine  18  are coupled by a second rotor shaft  32 .  
         [0013]     During operation, air flows axially through fan assembly  12 , in a direction that is substantially parallel to a central axis  34  extending through engine  10 , and compressed air is supplied to high pressure compressor  14 . The highly compressed air is delivered to combustor  16 . Combustion gas flow (not shown in  FIG. 1 ) from combustor  16  drives turbines  18  and  20 . Turbine  18  drives compressor  14  by way of shaft  32  and turbine  20  drives fan assembly  12  by way of shaft  31 .  
         [0014]      FIG. 2  is an enlarged sectional schematic illustration of a portion of gas turbine engine  10 . Combustor  16  includes, in the exemplary embodiment, an annular outer liner  40 , an annular inner liner  42 , and a domed end (not shown) extending between outer and inner liners  40  and  42 , respectively. Outer liner  40  and inner liner  42  are spaced radially inward from a combustor casing (not shown) and define a combustion chamber system assembly  46 . An inner nozzle support  44  is generally annular and extends downstream from a diffuser (not shown). Combustion chamber  46  is generally annular in shape and is defined between liners  40  and  42 . Inner liner  42  and inner nozzle support  44  define an inner passageway  50 . Outer and inner liners  40  and  42  each extend to a turbine nozzle  52  positioned downstream from combustor  16 .  
         [0015]     High pressure turbine  18  is coupled substantially coaxially with compressor  14  (shown in  FIG. 1 ) and downstream from combustor  16 . Turbine  18  includes a rotor assembly  54  that includes at least one rotor  56  that is formed by one or more disks  60 . In the exemplary embodiment, disk  60  includes an outer rim  62 , an inner hub (not shown), and an integral web  66  extending generally radially therebetween and radially inward from a respective blade dovetail slot  68 . Each disk  60  also includes a plurality of blades  70  extending radially outward from outer rim  62 . Circumscribing the row of high pressure blades  70  in close clearance relationship therewith is an annular shroud  72 . Shroud  72  may include a plurality of annular sectors attached at an inner side of an annular band  74  that is formed of a plurality of sectors that form a complete circle. Disk  60  extends circumferentially around rotor assembly  54  and each row of blades  70  is sometimes referred to as a turbine stage. Disk  60  includes an aft surface  80  and a fore surface  82 . An aft disk cavity  84  houses disk  60 .  
         [0016]     Stationary turbine nozzles  52  are located between combustor  16  and turbine blades  70 . Nozzles  52  direct the combustion gases toward turbine blades  70  and the impingement of the combustion gases on blades  70  impart a rotation of turbine disk  60 . A plurality of stationary stator vanes  86  direct the combustion gases passing through turbine blades  70  to the next turbine stage (not shown).  
         [0017]     A clearance control system  88  controls the clearance, or distance, between turbine blades  70  and turbine shroud  72 . Clearance control system  88  includes a cooling air supply pipe  90  connected at one end to an air supply source, for example, the middle stages of compressor  14 , or compressor  14  discharge, and at another end to a cooling air inlet  92  in a turbine casing  94 . A heat exchanger  96  is coupled to and is in flow communication with cooling air supply pipe  90 . Heat exchanger  96  includes a coolant loop  98  which removes heat from the cooling air as the air passes through heat exchanger  96 . Any suitable coolant can be used in coolant loop  98 . In one exemplary embodiment, the gas turbine engine fuel is directed through coolant loop  98 . In other exemplary embodiments, turbine bearing oil or a refrigerant is directed through coolant loop  98 .  
         [0018]     In operation, cooling air is directed through cooling air supply pipe  90  and heat exchanger  96  and into cooling air inlet  92  in turbine casing  94 . A coolant fluid flows through coolant loop  98  which removes heat from air passing through heat exchanger  96 . The cooling air flows into a cavity  100  between turbine casing  94  and turbine shroud  72  and is used to cool turbine casing  94  and turbine shroud  72 . The cooling air is then directed through stator vanes  86  and into disk cavity  84  where the cooling air is directed to aft surface  80  of disk  60  to cool disk  60 . In one exemplary embodiment, a portion of the cooling air is returned to cooling air supply pipe to be re-cooled in heat exchanger  96  and reused in the cooling of casing  94 , shroud  72  and disk  60 .  
         [0019]     While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.