Abstract:
A turbine airfoil for a gas turbine engine includes: (a) spaced-apart pressure and suction sidewalls extending between a leading edge and a trailing edge; (b) a first cavity disposed between the pressure and suction sidewalls, the first cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates with an exterior surface of the airfoil; (c) a second cavity disposed between the pressure and suction sidewalls, the second cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates solely with the suction sidewall of the airfoil; and (d) a metering structure adapted to substantially restrict air flow into the second cavity.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to gas turbine engine turbines and more particularly to methods for cooling turbine airfoils in such engines. 
         [0002]    A gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine (HPT) in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. 
         [0003]    The HPT includes annular arrays of stationary airfoils called vanes or nozzles that direct the gases exiting the combustor into rotating airfoils called blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. These components operate in an extremely high temperature environment, and must be cooled by air flow, typically impingement or film cooling, or a combination thereof, to ensure adequate service life. Typically, the air used for cooling is extracted from one or more points in the compressor. These bleed flows represent a loss of net work output and/or thrust to the thermodynamic cycle. They increase specific fuel consumption (SFC) and are generally to be avoided as much as possible. 
         [0004]    Typically, an HPT nozzle airfoil has a leading edge cavity and a trailing edge cavity separated by a rib or wall. The location of this wall is positioned to reduce the overall length of airfoil panels on each cavity, to avoid ballooning stresses. In addition, the position of the wall is dependent on the location of the inner band flange, relative to the leading edge cavity break out for casting producibility. As a result the wall between the two cavities is located at or near the throat area, which is the location of minimum cross-sectional area between two adjacent nozzle airfoils. Film holes, which are used to cool the suction side of the airfoil, are typically placed upstream of the throat area so as to make the flow non-chargeable to the engine cycle, avoiding a performance penalty. The film holes are placed as close to the throat as practical, to minimize the length of suction side surface dependent on this film for cooling. 
         [0005]    These suction side film holes discharge air into a lower pressure region of the gas path. The film hole cooling array and flow level is dependant on the pressure ratio from the supply cavity to the gas path discharge location. The supply pressure of the feed cavity is set to avoid ingestion anywhere across its wall, which is most likely to occur at the leading edge and pressure sides of the airfoil. As a result, the pressure ratio at the suction side film holes is excessively high. This results in a high flow rate per hole and a lower hole density within the array, effectively reducing cooling effectiveness. 
       BRIEF SUMMARY OF THE INVENTION 
       [0006]    These and other shortcomings of the prior art are addressed by the present invention, which provides a turbine airfoil with an internal cavity that is fed a reduced pressure cooling flow to improve film cooling effectiveness. 
         [0007]    According to one aspect, a turbine airfoil for a gas turbine engine includes: (a) spaced-apart pressure and suction sidewalls extending between a leading edge and a trailing edge; (b) a first cavity disposed between the pressure and suction sidewalls, the first cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates with an exterior surface of the airfoil; (c) a second cavity disposed between the pressure and suction sidewalls, the second cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates solely with the suction sidewall of the airfoil; and (d) a metering structure adapted to substantially restrict air flow into the second cavity. 
         [0008]    According to another aspect of the invention, a method is provided for, in a gas turbine engine, cooling a turbine nozzle having at least two spaced-apart, hollow, turbine airfoils, each of which includes: a first cavity disposed between pressure and suction sidewalls of the turbine airfoil and connected to at least one film cooling hole which communicates with an exterior surface of the airfoil; and a second cavity disposed between the pressure and suction sidewalls, and connected to at least one film cooling hole which communicates solely with the suction sidewall of the airfoil. The method includes: (a) directing cooling air from a source within the engine to each of the first cavities at a first pressure; (b) exhausting cooling air from the first cavities through the at least one film cooling hole connected thereto; (c) directing cooling air from a source within the engine to each of the second cavities; (d) dropping the pressure of the cooling air to a second pressure substantially lower than the first pressure before introducing it into each of the second cavities; and (e) exhausting cooling air from the second cavities through the at least one film cooling hole connected thereto. 
         [0009]    According to another aspect of the invention, a turbine airfoil for a gas turbine engine includes: (a) spaced-apart pressure and suction sidewalls extending between a leading edge and a trailing edge; (b) a first cavity disposed between the pressure and suction sidewalls, the first cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates with an exterior surface of the airfoil; (c) a second cavity disposed between the pressure and suction sidewalls, the second cavity being separated from the first cavity by a wall having at least one metering hole passing therethrough, and connected to at least one film cooling hole which communicates solely with the suction sidewall of the airfoil; and (d) a metering structure adapted to substantially restrict air flow into the second cavity. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]    The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
           [0011]      FIG. 1  a schematic cross-sectional view of a high-bypass gas turbine engine including a turbine nozzle constructed in accordance with the present invention; 
           [0012]      FIG. 2  is a perspective view of a turbine nozzle segment constructed in accordance with an aspect of the present invention; 
           [0013]      FIG. 3  is a view taken along lines  3 - 3  of  FIG. 2 ; 
           [0014]      FIG. 4  is another perspective view of the turbine nozzle shown in  FIG. 2 . 
           [0015]      FIG. 5  is a perspective view of an alternative turbine nozzle segment constructed in accordance with an aspect of the present invention; 
           [0016]      FIG. 6  is a view taken along lines  6 - 6  of  FIG. 5 ; and 
           [0017]      FIG. 7  is a another perspective view of the turbine nozzle shown in  FIG. 5 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0018]    Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  depicts a gas turbine engine  10  having a fan  12 , a low pressure compressor or “booster”  14  and a low pressure turbine (“LPT”)  16  collectively referred to as a “low pressure system”, and a high pressure compressor (“HPC”)  18 , a combustor  20 , and a high pressure turbine (“HPT”)  22 , collectively referred to as a “gas generator” or “core”. Together, the high and low pressure systems are operable in a known manner to generate a primary or core flow as well as a fan flow or bypass flow. While the illustrated engine  10  is a high-bypass turbofan engine, the principles described herein are equally applicable to turboprop, turbojet, and turboshaft engines, as well as turbine engines used for other vehicles or in stationary applications. 
         [0019]    The high pressure turbine  22  includes a high pressure nozzle  24 . As shown in  FIG. 2 , the high pressure nozzle  24  comprises an array of airfoil-shaped hollow vanes  26  that are supported between an arcuate, segmented inner band  28  and an arcuate, segmented outer band  30 . The vanes  26 , first inner band  28  and outer band  30  are arranged into a plurality of circumferentially adjoining nozzle segments  32  that collectively form a complete 360° assembly. In this example each of the nozzle segments  32  is a “singlet” having one vane  26 , but other configurations (doublet, triplet, etc.) as well as continuous rings or half-rings are known. The inner and outer bands  28  and  30  define the outer and inner radial flowpath boundaries, respectively, for the hot gas stream flowing through the high pressure nozzle  24 . The vanes  26  are configured so as to optimally direct the combustion gases to a rotor  33 . 
         [0020]    The rotor  33  includes an array of airfoil-shaped turbine blades  34  extending outwardly from a disk  36  that rotates about the centerline axis of the engine  10 . In the illustrated example, the high pressure turbine  22  is of the single-stage type having a single high pressure turbine nozzle  24  and rotor  26 . However, the principles of the present invention are equally applicable to multiple stage high-pressure turbines or to low-pressure turbines, where such turbines are cooled. 
         [0021]      FIGS. 3 and 4  illustrate the construction of the nozzle  24  in more detail. Each vane  26  has spaced-apart pressure and suction sidewalls  38  and  40  which extend between a leading edge  42  and a trailing edge  44 . The vanes  26  are arranged such that the suction sidewall  40  of a first vane  26  faces the pressure sidewall  38  of its neighboring vane  26 . The location at which the cross-sectional flow area between two neighboring vanes  26  is at a minimum is referred to as a “throat”, denoted “T” in  FIG. 3 . 
         [0022]    The interior of each vane  26  is generally hollow and is divided into a leading edge cavity  46  and a trailing edge cavity  48  by a rib or wall  50  integral to the vane casting. Optional impingement cooling inserts  52  and  54  of a known type pierced with impingement cooling holes  56  and  58  respectively are disposed in the leading and trailing edge cavities  46  and  48 , respectively. Film cooling holes  60  formed through the pressure sidewall  38  and leading edge  42  communicate with the leading and trailing edge cavities  46  and  48 . The leading and trailing edge cavities  46  and  48  may be fed cooling air from their radially inner or outer ends, or both. In this example the trailing edge cavity  48  has an inlet  62  at its radially outer end (see  FIG. 2 ), and the leading edge cavity  46  has an inlet  64  at its radially inner end (see  FIG. 4 ). Trailing edge cooling passages  66  such as the illustrated holes communicate with the aft end of the trailing edge cavity  48 . 
         [0023]    A metered cavity  68  is located aft of the leading edge cavity  46  and along the suction sidewall  40 . A plurality of film cooling holes  70  in the suction sidewall  40  communicate with the metered cavity  68 , and may have their exits located upstream of the throat T.  FIG. 3  is an example of a metered cavity  68  with a generally triangular cross-sectional shape ending just aft of the throat T. However, the shape and location of the metered cavity  68  is not critical and may be varied to suit a particular application. The metered cavity  68  may be fed from its radially inner or outer end, or both. As shown in  FIG. 2 , the metered cavity  68  is fed from its outer end. The radially outer end of the metered cavity  68  is closed off by a metering plate  72  with a metering hole  74  formed therethrough. The metering plate  72  is coupled to a source of cooling air, such as compressor discharge pressure (CDP) air, in a known manner. The metering hole  74  is sized to reduce the pressure in the metered cavity  68  to a selected level. 
         [0024]    In operation, pressurized cooling air is provided to the leading edge, trailing edge, and metered cavities,  46 ,  48 , and  68 . The cooling air passes into the leading edge and trailing edge cavities  46  and  48  at substantially the supply pressure. However, the cooling air flow supplied to the metered cavity  68  is restricted by the metering hole  74 , reducing pressure in the metered cavity  68  to a level just sufficient to provide positive film cooling of the suction sidewall  40  with acceptable backflow margin. This selected pressure level is substantially below the pressure in the leading edge and trailing edge cavities  46  and  48 . The resulting metered cavity pressure level enables the utilization of a higher density of the suction sidewall film cooling holes  70 , thereby providing more effective film cooling to the suction sidewall  40 . This cooling configuration provides effective cooling of the suction sidewall  40 , which historically exhibits thermal distress. The result is a more efficiently cooled airfoil while using substantially the same amount of cooling flow as the prior art. 
         [0025]      FIGS. 5-7  illustrate an alternative high pressure turbine nozzle  124 . It is generally similar in construction to the high pressure nozzle  24  described above and comprises an array of airfoil-shaped hollow vanes  126 , an arcuate, segmented inner band  128  and an arcuate, segmented outer band  130 . The vanes  126 , first inner band  128  and outer band  30  are arranged into a plurality of circumferentially adjoining “singlet” nozzle segments  132 . 
         [0026]      FIGS. 6 and 7  illustrate the construction of the nozzle  124  in more detail. Each vane  126  has spaced-apart pressure and suction sidewalls  138  and  140  which extend between a leading edge  142  and a trailing edge  144 . The vanes  126  are arranged such that the suction sidewall  140  of a first vane  126  faces the pressure sidewall  138  of its neighboring vane  126 . The location at which the cross-sectional flow area between two neighboring vanes  126  is at a minimum is referred to as a “throat”, denoted “T′” in  FIG. 6 . 
         [0027]    The interior of each vane  126  is generally hollow and is divided into a leading edge cavity  146  and a trailing edge cavity  148  by a rib or wall  150  integral to the vane casting. Optional impingement cooling inserts  152  and  154  of a known type pierced with impingement cooling holes  156  and  158  respectively are disposed in the leading and trailing edge cavities  146  and  148 , respectively. Film cooling holes  160  formed through the pressure sidewall  138  and leading edge  142  communicate with the leading and trailing edge cavities  146  and  148 . The leading and trailing edge cavities  146  and  148  may be fed cooling air from their radially inner or outer ends, or both. In this example the trailing edge cavity  148  has an inlet  162  at its radially outer end (see  FIG. 5 ), and the leading edge cavity  146  has an inlet  164  at its radially inner end (see  FIG. 7 ). Trailing edge cooling passages  166  such as the illustrated holes communicate with the aft end of the trailing edge cavity  148 . 
         [0028]    A metered cavity  168  is located aft of the leading edge cavity  146  and along the suction sidewall  140 . A plurality of film cooling holes  170  in the suction sidewall  140  communicate with the metered cavity  168 , and may have their exits located upstream of the throat T′.  FIG. 6  is an example of a metered cavity  168  defined by the wall  150  and another intersecting wall  151  and having a generally triangular cross-sectional shape ending just aft of the throat T′. The shape and location of the metered cavity  168  is not critical and may be varied to suit a particular application. The metered cavity  168  is feed by one or more metering holes  174  (only one of which is shown) formed in the intersecting wall  151 , which communicate with the trailing edge cavity  148 . Alternatively, the metering holes  174  could be formed through the wall  150  so as to feed the metered cavity  168  from the leading edge cavity  146 . The metering holes  174  are sized to reduce the pressure in the metered cavity  68  to a selected level. 
         [0029]    Operation of the turbine nozzle  124  is similar to that of the nozzle  24  described above. Pressurized cooling air is provided to the leading edge and trailing edge cavities  146  and  148 . The cooling air passes into the leading edge and trailing edge cavities  146  and  148  at substantially the supply pressure. Some of cooling air flow passes from the trailing edge cavity  148  through the metering hole  174 . The cooling air flow supplied to the metered cavity  168  is restricted by the metering hole  74 , reducing pressure in the metered cavity  168  to a level just sufficient to provide positive film cooling of the suction sidewall  140  with acceptable backflow margin. This selected pressure level is substantially below the pressure in the leading edge and trailing edge cavities  146  and  148 . The resulting metered cavity pressure level enables the utilization of a higher density of the suction sidewall film cooling holes  170 , thereby providing more effective film cooling to the suction sidewall  140 , as described above. 
         [0030]    The foregoing has described cooling arrangements for a gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation, the invention being defined by the claims.