Abstract:
A method of manufacturing a gas turbine engine includes providing a turbine mid-frame, coupling a plurality of rotor blades to a rotor disk, the rotor disk is coupled axially aft from the turbine mid-frame such that a cavity is defined between the rotor disk and the turbine mid-frame, and forming at least one opening extending through the turbine mid-frame to facilitate channeling cooling air into the gap, the opening configured to impart a high relative tangential velocity into the cooling air discharged from the opening.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engines and, more particularly, to methods and apparatuses for reducing turbine rotor temperatures. 
     Gas turbine engines typically include a compressor, a combustor, and a high-pressure turbine. In operation, air flows through the compressor and the compressed air is delivered to the combustor wherein the compressed air is mixed with fuel and ignited. The heated airflow is then channeled through the high-pressure turbine to facilitate driving the compressor. Moreover, during operation, un-cooled high-pressure turbine blades may transfer heat from the turbine blades, at gas path temperature, through the shank, and by conduction and/or convection, to the high-pressure turbine disk. Furthermore, cooling flow lost due to shank leaks my allow combustion gases to enter the cooling circuit, exposing the turbine disk to combustion gas temperatures. As a result, the turbine disk is exposed to high temperatures which may thermally fatigue the turbine disk. 
     To facilitate preventing damage that may result from turbine disk exposure to high temperatures and possibly combustion gases, at least one known gas turbine engine includes an internal cooling circuit to facilitate cooling the turbine disk. More specifically, cooling air is channeled along a forward face of the disk from a radially inner portion of the disk along a substantially linear path to a radially outer portion of the disk. However, channeling the cooling air linearly along the face of the rotor disk may not effectively cool the disk. Moreover, various fasteners and/or blade retainer pins within the cooling flowpath create undesired temperature rise due to windage, which may further reduce the ability for the cooling air to effectively cool the turbine disk. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, a method of manufacturing a gas turbine engine is provided. The method includes providing a turbine mid-frame, coupling a plurality of rotor blades to a rotor disk, the rotor disk is coupled axially aft from the turbine mid-frame such that a cavity is defined between the rotor disk and the turbine mid-frame, and forming at least one opening extending through the turbine mid-frame to facilitate channeling cooling air into the gap, the opening configured to impart a significant tangential velocity relative to the disk (swirl) in the cooling air discharged from the opening. 
     In another aspect, a turbine mid-frame assembly is provided. The turbine mid-frame assembly includes a turbine mid-frame including at least one of a fastener cover plate and an opening extending through the turbine mid-frame configured to facilitate cooling a turbine coupled downstream from and adjacent to the turbine mid-frame. 
     In a further aspect, a gas turbine engine is provided. The gas turbine engine includes a rotor disk, a plurality of blades coupled to the rotor disk, and a plurality of blade retaining devices coupled to an aft face of the rotor disk and the plurality of blades, the blade retaining devices configured to secure the plurality of blades to the rotor disk. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is schematic illustration of an exemplary gas turbine engine; 
         FIG. 2  is an enlarged cross-sectional view of a portion of the exemplary gas turbine engine shown in  FIG. 1 ; 
         FIG. 3  an enlarged view of a portion of the gas turbine engine rotor disk shown in  FIG. 2 ; 
         FIG. 4  is an end view of the gas turbine engine rotor disk shown in  FIG. 3 ; 
         FIG. 5  is a perspective view of an exemplary bolt cover; 
         FIG. 6  is an end view of the bolt cover shown in  FIG. 5 ; and 
         FIG. 7  is a cross-sectional view of a cooling opening shown in  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a schematic illustration of an exemplary gas turbine engine assembly  10  having a longitudinal axis  11 . Gas turbine engine assembly  10  includes a fan assembly  12 , a high-pressure radial compressor  14 , and a combustor  16 . Engine  10  also includes a high-pressure turbine assembly  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 . Fan assembly  12 , booster  22 , and low-pressure turbine  20  are coupled together by a first rotor shaft  32 , and compressor  14  and high-pressure turbine assembly  18  are coupled together by a second rotor shaft  34 . 
     In operation, air flows through fan assembly  12  and compressed air is supplied to high-pressure compressor  14  through booster  22 . The highly compressed air is delivered to combustor  16 . Hot products of combustion from combustor  16  are utilized to drive turbines  18  and  20 , which in turn drive fan assembly  12  and booster  22  utilizing first rotor shaft  32 , and also drive high-pressure compressor  14  utilizing second rotor shaft  34 , respectively. 
       FIG. 2  is an enlarged cross-sectional view of a portion of high-pressure turbine assembly  18  (shown in  FIG. 1 ).  FIG. 3  an enlarged cross-sectional view of a portion of high-pressure turbine rotor assembly  18  (shown in  FIG. 2 ).  FIG. 4  is an end view of a portion of high-pressure turbine rotor assembly  18  (shown in  FIG. 2 ). 
     In the exemplary embodiment, high-pressure turbine assembly  18  is coupled axially aft of a turbine mid-seal support structure  36  such that a cavity  38  is defined at least partially between mid-seal support structure  36  and high-pressure turbine assembly  18 . Gas turbine engine  10  also includes a mid-frame labyrinth seal  40  that is coupled to mid-seal support structure  36  to facilitate reducing and/or eliminating air and/or fluid from being channeled through an opening  42  defined between a radially inner portion of mid-seal support structure  36  and shaft  34  into cavity  38 . Moreover, gas turbine engine  10  includes a high-pressure turbine nozzle assembly  44  axially upstream from high-pressure turbine assembly  18  and a diffuser section  46 . In the exemplary embodiment, at least a portion of diffuser section  46 , high-pressure turbine nozzle assembly  44 , and mid-seal support structure  36  are coupled together using a plurality of mechanical fasteners  48 . In the exemplary embodiment, at last a portion of fastener  48 , i.e. a bolt head  50  extends at least partially into cavity  38 . 
     In the exemplary embodiment, high-pressure turbine assembly  18  includes a rotor disk  52  and a plurality of rotor blades  54  that are coupled to rotor disk  52 . Rotor blades  54  extend radially outward from rotor disk  52 , and each includes an airfoil  60 , a platform  62 , a shank  64 , and a dovetail  66 . Platform  62  extends between airfoil  60  and shank  64  such that each airfoil  60  extends radially outward from each respective platform  62 . Shank  64  extends radially inwardly from platform  62  to dovetail  66 . Dovetail  66  extends radially inwardly from shank  64  and facilitates securing each rotor blade  54  to rotor disk  52 . 
     Platform  62  includes an upstream side or skirt  70  and a downstream side or skirt  72 . Platform  62  also includes a forward angel wing  74 , and an aft angel wing  76  which each extend outwardly from respective skirts  70  and  72 . In the exemplary embodiment, each rotor blade  54  also includes a first portion  78  that extends radially inwardly from a lower surface  80  of aft angel wing  76  such that a first channel  82  is defined radially inwardly from each respective aft angel wing  76 . Moreover, rotor disk  52  includes a substantially L-shaped portion  84  that is coupled to an aft face  86  of rotor disk  52  such that a second channel  88  is defined radially outwardly from rotor disk  52 . In the exemplary embodiment, channel  82  is aligned substantially coaxially with channel  88  such that a cavity  90  is defined therebetween. In the exemplary embodiment, portion  84  is formed unitarily with rotor disk  52 . 
     High-pressure turbine rotor assembly  18  further includes a plurality of blade retaining devices  100  that are utilized to secure plurality of rotor blades  54  to rotor disk  52 . Each blade retaining device  100  has a width  102  that is selectively sized such that a radially outer edge  104  of blade retaining device  100  is positioned at least partially within channel  82  and a radially inner edge  106  of blade retaining device  100  is positioned at least partially within channel  88 . Moreover, each blade retaining device  100  has a length  108  that is sized to secure at least one rotor blade  54  to rotor disk  52 . In the exemplary embodiment, length  108  is selected to secure three rotor blades  54  to rotor disk  52 . Moreover, although the exemplary embodiment illustrates each blade retaining device  100  securing three rotor blades  54  to rotor disk  52 , it should be realized that length  108  can be selected to couple, one, two, three, or more rotor blades  54  to rotor disk  52 . 
     In the exemplary embodiment, blade retaining devices  100  are each fabricated from a flexible metallic material. During installation radially outer edge  104  is positioned within channel  82 , blade retaining device  100  is flexed and/or deformed such that radially inner edge  106  can be positioned within channel  88 . Blade retaining device  100  then returns to its normal or unflexed condition to facilitate maintaining blade retaining device  100  within channels  82  and  88 , respectively, and thus securing plurality of rotor blades  54  to rotor disk  52 . To facilitate cooling high-pressure turbine assembly  18 , gas turbine engine  10  further includes a bolt cover  120  and at least one opening  122  extending through turbine mid-seal support structure  36 . 
       FIG. 5  is a perspective view of bolt cover  120 .  FIG. 6  is an end view of bolt cover  120 . In the exemplary embodiment, bolt cover  120  includes a first side  130 , a second side  132  opposite first side  130 , and a radially inner portion  134  that is coupled between first and second sides  130  and  132 , respectively, Accordingly, and in the exemplary embodiment, bolt cover  120  has a substantially U-shaped cross-sectional profile. First side  130  includes a first quantity of slots  140  that are spaced circumferentially around a periphery of bolt cover  120 . Each slot  140  has a width  142  and a length  144  that are each selectively sized to at least partially circumscribe a respective bolt head  50 . More specifically, gas turbine engine  10  includes n bolts to facilitate coupling diffuser section  46 , high-pressure turbine nozzle assembly  44 , and mid-seal support structure  36  together. Accordingly, and in the exemplary embodiment, bolt cover  120  also includes n slots  140 , wherein each slot  140  at least partially circumscribes a respective bolt head  50 . In another embodiment, bolt cover  120  includes n-m slots  140 , wherein m is defined as a quantity of fasteners  48  that are utilized to couple bolt cover  120  to mid-seal support structure  36  as discussed herein. 
     Bolt cover second side  132  includes m openings  150  extending therethrough. Each opening  150  has a diameter  152  that is less than a diameter  154  of a respective bolt head  50 . In the exemplary embodiment, bolt cover  120  includes three openings  150 , i.e. m=3. In the exemplary embodiment, bolt cover  120  is coupled within gas turbine engine  10  to facilitate covering bolt heads  50  and thereby improve cooling flow within cavity  38 . 
     To install bolt cover  120 , bolt cover  120  is positioned within gas turbine engine  10  such that plurality of slots  140  each at least partially circumscribe a respective bolt head  50 . More specifically, slots  140  are selectively sized such that bolt cover  120  can be installed within gas turbine engine  10  without removing all of the fasteners  48 . Accordingly, only m fasteners are removed and/or not installed. The m fasteners  48  are then inserted through respective openings  150  to facilitate coupling bolt cover  120  within gas turbine engine  10 . Since each opening  150  is smaller than a respective bolt head  50 , coupling a nut  160  to a respective fastener  48  facilitates securing bolt cover  120  within cavity  38 . Since bolt cover  120  has a substantially U-shaped cross-sectional profile, bolt heads  50  are positioned within a cavity  162  that is defined between first side  130  and second side  132 . Moreover, second side  132  facilitates channeling air around bolt heads  50  and thus facilitate reducing air turbulence within cavity  38  that would be-created with exposed bolt heads extending into cavity  38 . 
     To facilitate cooling high-pressure turbine assembly  18 , gas turbine engine  10  includes a plurality of openings  122  extending through turbine mid-seal support structure  36 . More specifically, openings  122  extend through turbine mid-seal support structure  36  and into flow communication with cavity  38 . 
     More specifically, and as shown in  FIG. 7 , each opening  122  includes an axially component  190  and a tangential component  192  such that a high relative tangential velocity is induced into cooling air  194  channeled through each opening  122 . Swirl, as used herein, is defined as a ratio of the tangential cooling air velocity to the velocity of rotating high-pressure turbine assembly  18 . More specifically, opening  122  facilitates increasing a velocity of cooling air  194  channeled through opening  122  to a velocity that is greater than the velocity of high-pressure turbine assembly  18  during operation. 
     In one embodiment, opening  122  is formed through turbine mid-seal support structure  36  at a tangential angle between approximately forty-five degrees and approximately 80 degrees with respect to centerline axis  11 . In the exemplary embodiment, opening  122  is formed through turbine mid-seal support structure  36  at a tangential angle that is approximately seventy degrees with respect to centerline axis  11 . 
     During operation, cooling air  194  is channeled through openings  122  to facilitate cooling high-pressure turbine assembly  18 . More specifically, cooling air  194  is channeled through openings  122  an angle that is tangent to high-pressure turbine assembly  18  such that swirl is induced into cooling air  194 . Cooling air  194  is then channeled over an exterior surface of bolt cover  120  which facilitates reducing and/or eliminating drag induced temperature rise (windage) that may be introduced into the cooling air caused by bolt heads  50 . Additionally, blade retaining devices  100  facilitate reducing and/or eliminating airflow leakage through high-pressure turbine assembly  18  by substantially sealing any gaps that may exist between dovetail  66  and rotor  52 . 
     The above-described high-pressure turbine rotor cooling system is cost-effective and highly reliable. The cooling system includes at least one opening to facilitate channeling cooling air into a cavity that is between the turbine mid-frame support and the high-pressure turbine rotor. The opening is formed such that the a swirling motion is imparted to the cooling air channeled therethrough. Moreover, the cooling system described herein includes a bolt cover to facilitate reducing turbulence within the cavity, and a plurality of blade retaining devices that are utilized to secure the rotor blades to the rotor disk and also to facilitate reducing and/or eliminating any airflow leakage that may occur between the turbine blades and the turbine rotor. As a result, the cooling air channeled into the cavity more effectively cools the high pressure turbine rotor compared to known cooling methods to facilitate extending a useful life of the rotor blades in a cost-effective and reliable manner. 
     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.