Abstract:
Methods and apparatus for cooling gas turbine rotor blades is provided. The blade includes an airfoil having an internal three pass serpentine cooling circuit having 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 radially extending internal rib. The serpentine cooling circuit includes a first inlet in flow communication with the first cavity and a second inlet in flow communication with at least one of the second and third cavities wherein the first and second inlets are formed during casting of the airfoil.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines and more particularly, to methods and apparatus for cooling gas turbine engine rotor assemblies. 
   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. 
   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. To improve cooling efficiency a refresher hole is drilled through the root to permit new flow of the gas to enter the blade and intersect the root turn of a serpentine passage. The refresher holes are of a relatively small diameter such that the gases passes through the holes are raised in temperature due to high velocity. Refresher holes are sized to a relatively small diameter to meter the amount of mixed gas. Drilling the holes adds an extra operation during the manufacturing process that is expensive and labor intensive. 
   BRIEF DESCRIPTION OF THE INVENTION 
   In one embodiment, a gas turbine rotor blade includes an airfoil having an internal three pass serpentine cooling circuit having 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 radially extending internal rib. The serpentine cooling circuit includes a first inlet in flow communication with the first cavity and a second inlet in flow communication with at least one of the second and third cavities wherein the first and second inlets are formed during casting of the airfoil. 
   In another embodiment, a method for cooling a gas turbine engine turbine blade wherein the turbine blade includes a serpentine cooling circuit extending between a dovetail of the blade and a tip of the blade and a flow metering device coupled to the dovetail. The method includes providing a first flow of a cooling gas to the blade through a first cooling inlet, providing a second flow of a cooling gas to the blade through a second cooling inlet, and controlling the cooling gas flow through the first and second cooling inlets using the flow metering device. 
   In yet another embodiment, a gas turbine engine assembly includes a compressor, a combustor, and a turbine coupled to the compressor. The turbine includes an airfoil having an internal three pass serpentine cooling circuit having 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 radially extending internal rib. The serpentine cooling circuit including a first inlet in flow communication with the first cavity and a second inlet in flow communication with at least one of the second and third cavities wherein the first and second inlets are formed during casting of the airfoil. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a schematic illustration of an exemplary gas turbine engine; 
       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 
       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 
     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 . 
   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 . 
     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. 
   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 . 
   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 tip plate  54  which 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. An inlet passage  64  is configured to channel air into first cavity  58  and then 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 . 
   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  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 . The cooler air reduces the temperature and increases the pressure of the 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 . However, the velocity of the air passing through hole  66  is relativity high causing the air temperature of the air entering third cavity  62  to be higher than the temperature of the air entering hole  66  such that a cooling efficiency of the refresher air is less than optimal. 
     FIG. 3  is a perspective internal schematic illustration of a rotor blade  300  in accordance with an exemplary embodiment of the present invention. Blade  300  includes a hollow airfoil  302  and an integral dovetail  304  used for mounting airfoil  302  to a rotor disk (not shown). 
   Airfoil  302  includes a first sidewall  306  (shown cutaway) and a second sidewall  308 . First sidewall  306  is convex and defines a suction side of airfoil  302 , and second sidewall  308  is concave and defines a pressure side of airfoil  302 . Sidewalls  306  and  308  are connected at a leading edge  310  and at an axially-spaced trailing edge  312  of airfoil  302  that is downstream from leading edge  310 . 
   First and second sidewalls  306  and  308 , respectively, extend longitudinally or radially outward to span from a blade root  314  positioned adjacent dovetail  44  to a tip plate  316  which defines a radially outer boundary of an internal cooling chamber  318 . Cooling chamber  318  is defined within airfoil  302  between sidewalls  306  and  308 . In the exemplary embodiment, cooling chamber  318  includes a serpentine passage comprising a first cavity  320 , a second cavity  322  and a third cavity  324  cooled with compressor bleed air. An inlet passage  326  is configured to channel air into first cavity  320  and then into second cavity  322 . A refresher inlet  328  couples second cavity  322  to the compressor bleed air. A downstream end of third cavity  324  is in flow communication with a plurality of trailing edge slots  332  which extend longitudinally (axially) along trailing edge  312 . Particularly, trailing edge slots  332  extend along pressure side wall  308  to trailing edge  312 . 
   In the exemplary embodiment, both inlet  326  and refresher inlet  328  are formed during the casting process of blade  300 . The ceramic core used to cast blade  300  includes a tab that extends through the passages where inlet  326  and refresher inlet  328  are formed. The tabs are used to secure the core in the casting mold. Using two tabs permits a more secure connection than is available using only one tab through the inlet passage in the prior art blade. Additionally, trailing edge slots  332  are also cast rather than drilled, as in the prior art blade. Each of the slots also include a tab extending from the casting mold and are used to further secure the ceramic core during casting. 
   In operation, cooling air is supplied to blade  300  from compressor bleed air through inlet  326  and refresher inlet  328 . Air entering blade  300  through inlet  326  is directed through first cavity  320  and into second cavity  322 . Refresher inlet  328  permits cooler compressor bleed air to enter chamber  318  between second cavity  322  and third cavity  324 . The cooler air reduces the temperature and increases the pressure of the air entering third cavity  324 . The cooler air and increased pressure facilitate cooling trailing edge  312  through slots  332 . Air entering first cavity  320  is metered using a metering plate  330 , which includes a first hole  333  of a predetermined size. Metering plate  330  includes a second hole  334  of a predetermined size to control the floe of refresher air through refresher inlet  328 . The flow and pressure in cooling chamber  318  is adjusted by grinding metering plate  330  from dovetail  44  and installing a new metering plate  330  with a different diameter holes  333  and  334 . The velocity of the air passing through refresher inlet  328  is reduced with respect to the velocity of the air passing through hole  66  of the prior art blade  40 . Because the velocity is less the temperature raise of the air entering third cavity  324  is less such that a cooling efficiency of the refresher air is facilitated being optimal. 
   The above-described cast refresher flow passage is a cost-effective and highly reliable method for reducing a temperature rise of the cooling air due to lower Mach numbers resulting in improved airfoil cooling efficiency, eliminating stress concentration from sharp edge of the machined refresher hole and eliminating the recast layer/micro-cracks associated with EDM, improved core support during the casting process, eliminating long EDM hole machining. The above described method also produces less scrap. The prior art design blade is scrapped if the refresher hole is oversized. A blade fabricated using the method described above is corrected for an oversized refresher hole by grinding off the metering plate and replacing it with a new one. Further, flow splits between the inlet and refresher inlet can be adjusted using the metering plate holes. Accordingly, the cast refresher flow passage assembly facilitates operating gas turbine engine components, 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.