Abstract:
Methods and apparatus for cooling gas turbine rotor blades is provided. The rotor blades include an airfoil having a pressure sidewall and a second suction sidewall connected together at a leading edge and a trailing edge, such that an internal three pass serpentine cooling circuit is formed therebetween. The cooling circuit includes radially extending first, second, and third serpentine cooling cavities partially separated by, in axially aft succession, a first radially extending internal rib and a second internal rib. The second rib includes a radially inner first portion and a radially outer portion wherein the radially outer portion is angled obliquely with respect to the first portion.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to gas turbine engines and more particularly, to methods and apparatus for cooling gas turbine engine rotor assemblies.  
         [0002]     Turbine rotor assemblies typically include at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side, and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. Known blades are hollow such that an internal cooling cavity is defined at least partially by the airfoil, platform, shank, and dovetail.  
         [0003]     At least some known high pressure turbine blades include an internal cooling cavity that is serpentine such that a path of cooling gas is channeled radially outward to the blade tip where the flow reverses direction and flows back radially inwardly toward the blade root. The flow may exit the blade through the root or the flow may be directed to holes in the trailing edge to permit the gas to flow across a surface of the trailing edge for cooling the trailing edge. In cooled turbine blades, the internal pressure of cooing air is attempted to be maintained greater than the local external pressure in the area of the blade. The amount by which the internal pressure exceeds the external pressure is typically referred to as positive Back Flow Margin (BFM). Having a positive BFM prevents hot gas ingestion into the blade interior in the event of a breached wall or severe cycle deterioration.  
         [0004]     Furthermore, the aft tip region typically operates at an elevated temperature with respect to the rest of the blade such that film cooling in this area is desirable to improve blade life. In some known blades this film cooling is provided by using film holes in flow communication with a third or aftmost cavity in the cooling circuit. However, adequate internal pressure in the third cavity may not be able to be maintained in all cases. The second cavity or the cavity adjacent and upstream of the third cavity has adequate pressure but is located too far forward to be able to provide film cooling where it is needed.  
       BRIEF DESCRIPTION OF THE INVENTION  
       [0005]     In one embodiment, a gas turbine rotor blade includes an airfoil having a pressure sidewall and a second suction sidewall connected together at a leading edge and a trailing edge, such that an internal three pass serpentine cooling circuit is formed therebetween. The cooling circuit includes radially extending first, second, and third serpentine cooling cavities partially separated by, in axially aft succession, a first radially extending internal rib and a second internal rib. The second rib includes a radially inner first portion and a radially outer portion wherein the radially outer portion is angled obliquely with respect to the first portion.  
         [0006]     In another embodiment, a method for cooling a gas turbine engine turbine blade is provided. The turbine blade includes an airfoil having a pressure sidewall and a suction sidewall connected together at a leading edge and a trailing edge, and a cooling circuit including radially extending first, second, and third serpentine cooling cavities partially separated by, in axially aft succession, a first radially extending internal rib and a second internal rib such that an internal three pass serpentine cooling circuit is formed that extends between a dovetail of the blade and a tip of the blade. The second rib includes a radially inner first portion and a radially outer portion wherein the radially outer portion is angled obliquely with respect to the first portion. The method includes providing a flow of a cooling gas to the blade through a cooling gas inlet, channeling the flow of the cooling gas through the first cavity using the first rib, channeling the flow of the cooling gas into the second cavity using the second rib, and directing at least a portion of the flow of the cooling gas through at least one film hole communicatively coupled between the second cavity and an external surface of the pressure sidewall.  
         [0007]     In yet another embodiment, a gas turbine engine assembly includes a compressor, a combustor, and a turbine coupled to the compressor the turbine including a rotor blade that includes an airfoil having a pressure sidewall and a suction sidewall connected together at a leading edge and a trailing edge, such that an internal three pass serpentine cooling circuit is formed therebetween, the cooling circuit including radially extending first, second, and third serpentine cooling cavities partially separated by, in axially aft succession, a first radially extending internal rib and a second internal rib. The second rib includes a radially inner first portion and a radially outer portion wherein the radially outer portion is angled obliquely with respect to the first portion. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0008]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine;  
         [0009]      FIG. 2  is a perspective internal schematic illustration of a known rotor blade that may be used with the gas turbine engine shown in  FIG. 1 ; and  
         [0010]      FIG. 3  is a perspective internal schematic illustration of a rotor blade in accordance with an exemplary embodiment of the present invention. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0011]      FIG. 1  is a schematic cross-sectional illustration of a gas turbine engine  10  including an inlet  12 , an inlet particle separator  14 , core inlet guide vanes  16 . Engine  10  also includes in serial flow communication an axial compressor  18 , a radial compressor  20  or impellor, and a deswirler diffuser  22 . Downstream from deswirler diffuser  22  is a combustor  24 , a high pressure turbine  26  and a power turbine  28 .  
         [0012]     In operation, air flows through inlet  12  to axial compressor  18  and to radial compressor  20 . The highly compressed air is delivered to combustor  24 . The combustion exit gases are delivered from combustor  24  to high pressure turbine  26  and power turbine  28 . Flow from combustor  24  drives high pressure turbine  26  and power turbine  28  coupled to a rotatable main turbine shaft  30  aligned with a longitudinal axis  32  of gas turbine engine  10  in an axial direction and exits gas turbine engine  10  through an exhaust system  34 .  
         [0013]      FIG. 2  is a perspective internal schematic illustration of a known rotor blade  40  that may be used with gas turbine engine  10  (shown in  FIG. 1 ). In an exemplary embodiment, a plurality of rotor blades  40  form a high pressure turbine rotor blade stage (not shown) of gas turbine engine  10 . Each rotor blade  40  includes a hollow airfoil  42  and an integral dovetail  44  used for mounting airfoil  42  to a rotor disk (not shown) in a known manner.  
         [0014]     Airfoil  42  includes a first sidewall  45  (shown cutaway) and a second sidewall  46 . First sidewall  45  is convex and defines a suction side of airfoil  42 , and second sidewall  46  is concave and defines a pressure side of airfoil  42 . Sidewalls  45  and  46  are connected at a leading edge  48  and at an axially-spaced trailing edge  50  of airfoil  42  that is downstream from leading edge  48 .  
         [0015]     First and second sidewalls  45  and  46 , respectively, extend longitudinally or radially outward to span from a blade root  52  positioned adjacent dovetail  44  to a squeeler tip  53  comprising a tip plate  54  that recessed with respect to a blade end  55 . Tip plate  54  defines a radially outer boundary of an internal cooling chamber  56 . Cooling chamber  56  is defined within airfoil  42  between sidewalls  45  and  46 . In the exemplary embodiment, cooling chamber  56  includes a serpentine passage comprising a first cavity  58 , a second cavity  60  and a third cavity  62  cooled with compressor bleed air. First cavity  58  and second cavity  60  are separated by a first rib  63  extending radially outward from root  52  towards tip  54 . A second rib  65  extends radially inward from tip  54  towards root  52  and spaced axially downstream from rib  63 . Second rib  65  separates cavity  60  from cavity  62 . An inlet passage  64  is configured to channel air into first cavity  58  and then around first rib  63  into second cavity  60 . A refresher hole  66  couples second cavity  60  to the compressor bleed air. Refresher hole  66  is formed using an electrical discharge machining (EDM) process that generates stress concentration at the sharp edge surrounding the openings of refresher hole  66  and generates recast layer/micro-cracks associated with the EDM process. A downstream end of third cavity  62  is in flow communication with a plurality of trailing edge holes  70  which extend longitudinally (axially) along trailing edge  50 . Particularly, trailing edge holes  70  extend along pressure side wall  46  to trailing edge  50 .  
         [0016]     In operation, cooling air is supplied to blade  40  from compressor bleed air through inlet  64  and refresher hole  66 . Air entering blade  40  through inlet  64  is directed through first cavity  58 , a round rib  63  and into second cavity  60 . Refresher hole  66  permits cooler compressor bleed air to enter chamber  56  between second cavity  60  and third cavity  62  proximate a radially inner end  76  of rib  65 . The cooler air entering from refresher hole  66  facilitates reducing the temperature and increasing the pressure of the cooling air entering third cavity  62 . The cooler air and increased pressure facilitate cooling trailing edge  50  through holes  70 . Air entering first cavity  58  is metered using a meter plate  68 , which includes a hole  69  of a predetermined size. The flow and pressure in first cavity  58  is adjusted by grinding metering plate  68  from dovetail  44  and installing a new metering plate  68  with a different diameter hole  69 . The flow and pressure in third cavity  62  is adjusted by modifying the size of hole  66 .  
         [0017]     During fabrication of blade  40 , a casting core (not shown) is used to form the shape of blade  40  inside a mold. The casting core includes a relatively large tip support in third cavity  62 . Accordingly, a relatively large area tip hole  80  is used to remove the core after casting. Tip hole  80  tends to reduce the back flow margin in third cavity  62  such that adding film holes to aid film cooling of the blade tip may result in a low pressure feeding the film holes from third cavity  62 . Such low pressure may lead to hot gas ingestion causing additional distress to the blade tip.  
         [0018]      FIG. 3  is a perspective internal schematic illustration of a rotor blade  340  in accordance with an exemplary embodiment of the present invention. In an exemplary embodiment, cast pressure side cooling slots are used for core support during fabrication such that tip core support hole  80  is eliminated and the internal rib between second cavity and third cavity is curved towards the third cavity such that film cooling holes are supplied cooling air from the second cavity to maintain a higher internal pressure for a majority of the blade tip.  
         [0019]     Airfoil  342  includes a first sidewall  345  (shown cutaway) and a second sidewall  346 . First sidewall  345  is convex and defines a suction side of airfoil  342 , and second sidewall  346  is concave and defines a pressure side of airfoil  342 . Sidewalls  345  and  346  are connected at a leading edge  348  and at an axially-spaced trailing edge  350  of airfoil  342  that is downstream from leading edge  348 .  
         [0020]     First and second sidewalls  345  and  346 , respectively, extend longitudinally or radially outward to span from a blade root  352  positioned adjacent dovetail  344  to a squeeler tip  353  comprising a tip plate  54  that is recessed with respect to a blade end  355 . Tip plate  354  defines a radially outer boundary of an internal cooling chamber  356 . Cooling chamber  356  is defined within airfoil  342  between sidewalls  345  and  346 . In the exemplary embodiment, cooling chamber  356  includes a serpentine passage comprising a first cavity  358 , a second cavity  360  and a third cavity  362  cooled with compressor bleed air. First cavity  358  and second cavity  360  are separated by a first rib  363  extending radially outward from root  352  towards tip  354 . A second rib  365  extends radially inward from tip  354  towards root  352  and spaced axially downstream from rib  363 . Second rib  365  separates cavity  360  from cavity  362 . A radially out end  376  of rib  365  is curved towards cavity  362  such that end  376  intersects tip plate  354  farther aft or downstream than rib  65  intersects tip plate  54  (shown in  FIG. 2 ). One or more tip film holes  380  extend through sidewall  346  to permit cooling air from cavity  360  to exit blade  340  and form a cooling film at a blade end  355 . Tip film holes  380  extend through sidewall  346  from a point radially inward from tip plate  354  to an exit point on sidewall  345  that is radially outward from tip plate  354 . An inlet passage  364  is configured to channel air into first cavity  358 , around rib  363  and then into second cavity  360 . A refresher hole  366  couples second cavity  360  to compressor discharge air. Refresher hole  366  is formed using an electrical discharge machining (EDM) process. A downstream end of third cavity  362  is in flow communication with a plurality of trailing edge holes  370  which extend longitudinally (axially) along trailing edge  350 . Particularly, trailing edge holes  370  extend along pressure side wall  346  to trailing edge  350 .  
         [0021]     In operation, cooling air is supplied to blade  340  from compressor discharge air through inlet  364  and refresher hole  366 . Air entering blade  340  through inlet  364  is directed through first cavity  358 , around rib  363 , and into second cavity  360 . A portion of the air entering cavity  360  is channeled out of blade  340  through holes  380 . The exited air forms a film of relatively cool air at tip  382  and the film extends from sidewall  346 , over tip  382  and onto sidewall  345  such that a radially outer portion of sidewall  346 , a portion of tip  382 , and a portion of a radially outer portion of sidewall  345  is facilitated being cooled using the film. Curving end  376  permits locating holes  380  in a position such that the film formed over tip  382  provides a predetermined amount of cooling to tip  382 . Additionally, providing air at the entrance of cavity  360  to form the film improves BFM and cooling efficiency.  
         [0022]     Refresher hole  366  permits compressor discharge air that is cooler than the air in cavity  360  to enter chamber  356  between second cavity  360  and third cavity  362 . The cooler air reduces the temperature and increases the pressure of the air entering third cavity  362 . The cooler air and increased pressure facilitate cooling trailing edge  350  through holes  370 . Air entering first cavity  358  is metered using a meter plate  368 , which includes a hole  369  of a predetermined size. The flow and pressure in first cavity  358  is adjusted by grinding metering plate  368  from dovetail  344  and installing a new metering plate  368  with a different diameter hole  369 . The flow and pressure in third cavity  362  is adjusted by modifying the size of hole  366 . However, the velocity of the air passing through hole  366  is relativity high causing the air temperature of the air entering third cavity  362  to be higher than the temperature of the air entering hole  366  such that a cooling efficiency of the refresher air is less than optimal.  
         [0023]     The above-described internal aft curved rib is a cost-effective and highly reliable method for providing a source of film cooling air the blade aft tip region that is higher in pressure and lower in temperature than prior art blades. Accordingly, the internal aft curved rib facilitates operating gas turbine engine components, in a cost-effective and reliable manner.  
         [0024]     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.