Abstract:
A turbine nozzle for a gas turbine engine includes: (a) spaced-apart arcuate inner and outer bands; (b) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the interior of the vane defining at least a forward cavity and a mid-cavity positioned aft of the forward cavity; (c) a hollow impingement insert received inside the mid-cavity, the impingement insert having walls which are pierced with at least one impingement cooling hole; (d) a passage in the turbine vane at a radially outer end of the forward cavity adapted to be coupled to a source of cooling air; and (e) a passage in the inner band in fluid communication with a radially inner end of the forward cavity and a radially inner end of the impingement insert.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engine turbines and more particularly to methods for cooling turbine nozzles of such engines. 
     A gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure or gas generator turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. In a turbojet or turbofan engine, the core exhaust gas is directed through an exhaust nozzle to generate thrust. A turboshaft engine uses a low pressure or “work” turbine downstream of the core to extract energy from the primary flow to drive a shaft or other mechanical load. 
     The gas generator turbine includes annular arrays of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship. These components operate in an extremely high temperature environment, and must be cooled by air flow 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. 
     Various methods are known for cooling turbine components including film cooling, internal convection, and impingement. Impingement is known to be a particularly effective cooling method and is frequently used in large turbine engines where the engine core flow is substantial. However, higher turbine stages in small turboshaft and turboprop engines stage do not typically employ impingement cooling of the airfoil because there is either not enough cooling air or enough supply pressure available. Instead internal features like turbulators or pins provide the necessary convection heat transfer enhancements. 
     BRIEF SUMMARY OF THE INVENTION 
     These and other shortcomings of the prior art are addressed by the present invention, which provides a turbine nozzle cooled with a combination of impingement and convection cooling. 
     According to one aspect of the invention, a turbine nozzle for a gas turbine engine includes: (a) spaced-apart arcuate inner and outer bands; (b) a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the interior of the vane defining at least a forward cavity and a mid-cavity positioned aft of the forward cavity; (c) a hollow impingement insert received inside the mid-cavity, the impingement insert having walls which are pierced with at least one impingement cooling hole; (d) a passage in the turbine vane at a radially outer end of the forward cavity adapted to be coupled to a source of cooling air; and (e) a passage in the inner band in fluid communication with a radially inner end of the forward cavity and a radially inner end of the impingement insert. 
     According to another aspect of the invention, a method is provided for cooling a turbine nozzle of a gas turbine engine which includes: spaced-apart arcuate inner and outer bands; a hollow, airfoil-shaped turbine vane extending between the inner and outer bands, the interior of the vane defining at least a forward cavity and a mid-cavity positioned aft of the forward cavity; and a hollow impingement insert received inside the mid-cavity, the impingement insert having walls which are pierced with at least one impingement cooling hole. The method includes: (a) supplying cooling air to the forward cavity at a radially outer end thereof; (b) subsequently passing at least a portion of the cooling air entering the forward cavity from a radially inner end of the forward cavity to the radially inner end of the impingement insert; and (c) ejecting cooling air through the impingement cooling holes to cool the mid-cavity. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
         FIG. 1  is a schematic cross-sectional view of a turbine section constructed in accordance with the present invention; 
         FIG. 2  is an exploded perspective view of a turbine nozzle shown in  FIG. 1 ; 
         FIG. 3  is a perspective view of an impingement insert shown in  FIG. 2 ; and 
         FIG. 4  is a cutaway view of an assembled turbine nozzle and insert. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  depicts a portion of a gas generator turbine  10 , which is part of a gas turbine engine of a known type. The function of the gas generator turbine  10  is to extract energy from high-temperature, pressurized combustion gases from an upstream combustor (not shown) and to convert the energy to mechanical work, in a known manner. The gas generator turbine  10  drives an upstream compressor (not shown) through a shaft so as to supply pressurized air to the combustor. 
     In the illustrated example, the engine is a turboshaft engine and a work turbine would be located downstream of the gas generator turbine  10  and coupled to an output shaft. However, the principles described herein are equally applicable to turboprop, turbojet, and turbofan engines, as well as turbine engines used for other vehicles or in stationary applications. 
     The gas generator turbine  10  includes a first stage nozzle  12  which comprises a plurality of circumferentially spaced airfoil-shaped hollow first stage vanes  14  that are supported between an arcuate, segmented first stage outer band  16  and an arcuate, segmented first stage inner band  18 . The first stage vanes  14 , first stage outer band  16  and first stage inner band  18  are arranged into a plurality of circumferentially adjoining nozzle segments that collectively form a complete 360° assembly. The first stage outer and inner bands  16  and  18  define the outer and inner radial flowpath boundaries, respectively, for the hot gas stream flowing through the first stage nozzle  12 . The first stage vanes  14  are configured so as to optimally direct the combustion gases to a first stage rotor  20 . 
     The first stage rotor  20  includes a array of airfoil-shaped first stage turbine blades  22  extending outwardly from a first stage disk  24  that rotates about the centerline axis of the engine. A segmented, arcuate first stage shroud  26  is arranged so as to closely surround the first stage turbine blades  22  and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the first stage rotor  20 . 
     A second stage nozzle  28  is positioned downstream of the first stage rotor  20 , and comprises a plurality of circumferentially spaced airfoil-shaped hollow second stage vanes  30  that are supported between an arcuate, segmented second stage outer band  32  and an arcuate, segmented second stage inner band  34 . The second stage vanes  30 , second stage outer band  32  and second stage inner band  34  are arranged into a plurality of circumferentially adjoining nozzle segments  36  (see  FIG. 2 ) that collectively form a complete 360° assembly. The second stage outer and inner bands  32  and  34  define the outer and inner radial flowpath boundaries, respectively, for the hot gas stream flowing through the second stage turbine nozzle  34 . The second stage vanes  30  are configured so as to optimally direct the combustion gases to a second stage rotor  38 . 
     The second stage rotor  38  includes a radial array of airfoil-shaped second stage turbine blades  40  extending radially outwardly from a second stage disk  42  that rotates about the centerline axis of the engine. A segmented arcuate second stage shroud  44  is arranged so as to closely surround the second stage turbine blades  40  and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the second stage rotor  38 . 
     The segments of the first stage shroud  26  are supported by an array of arcuate first stage shroud hangers  46  that are in turn carried by an arcuate shroud support  48 , for example using the illustrated hooks, rails, and C-clips in a known manner. The second stage nozzle  28  is supported in part by mechanical connections to the first stage shroud hangers  46  and the shroud support  48 . 
       FIGS. 2-4  illustrate the construction of the second stage nozzle  28  in more detail.  FIG. 2  shows an individual nozzle segment  36  which is a “singlet” casting. It incorporates a segment  50  of the outer band  32 , a segment  52  of the inner band  34 , and a hollow second stage vane  30 . The interior of the second stage vane  30  is divided into a forward cavity  54 , a mid-cavity  56 , and a rear cavity  58 . An impingement insert  60  is received in the mid-cavity  56 . The radially outer end of the second stage vane  30  is closed by a cover  62 . The cover  62  is a plate-like structure which has a lower peripheral edge  64  that mates with an opening  66  at the radially outer end of the second stage vane  30 . 
       FIG. 3  shows the impingement insert  60  in more detail. The impingement insert  60  is a hollow, roughly airfoil-shaped structure, and has pressure and suction side walls  68  and  70  that extend between a leading edge  72  and a trailing edge  74 . The impingement insert  60  is closed off by a tip wall  76  (see  FIG. 2 ) and a root wall  78 . The walls of the impingement insert  60  are perforated by a plurality of impingement holes  80  of a known type which are sized to direct impingement jets against the walls of the mid-cavity  56 . An inlet tube  82  with a closed distal end protrudes radially inward from the root wall  78 . An inlet hole  84  is formed in the sidewall of the inlet tube  82 . 
       FIG. 4  shows in more detail how the impingement insert  60  is mounted in the second stage vane  30 . The impingement insert  60  is received in the mid-cavity  56 . Because the impingement insert  60  and the mid-cavity  56  both have larger cross-sectional areas at their outer ends than at their inner ends, the impingement insert  60  is installed from the outer end of the second stage vane  30 . Its inner end is secured to the inner band segment  52 , for example by a braze joint between the inlet tube  82  and the opening  87  in the inner band segment  52 . Welding or mechanical fasteners could also be used in place of the brazed joint, with some means of sealing. The outer end of the impingement insert  60  is free to move radially in or out as a result of thermal expansion or contraction during operation. One or more pads or protrusions (not shown) are provided as part of, or attached to, the impingement insert  60  and/or the walls of the mid-cavity  56  in order to locate the impingement insert and restrain its motion in the lateral and fore-and-aft directions. 
     Securing the impingement insert  60  at the inner end in this manner will effectively seal the cooling air entrance to the impingement insert  60  while allowing the impingement insert  60  to be installed from the radially outer end of the second stage vane  30 . This configuration allows maximum cooling air flow to be used in cooling the forward cavity  54  of the second stage vane  20 , and then utilizes the most effective convective cooling method to cool the mid-cavity  56  with the smallest amount of cooling flow possible. 
     In operation, compressor discharge air (CDP), at the highest pressure in the compressor, or another suitable cooling air flow, is ducted to a passage  83  at the radially outer end of the second stage vane  30 . It then flows radially inward through the forward cavity  54  where it cools the turbine vane  30  by convection. Although not shown, heat-transfer-enhancing structures such as fins, pins, turbulence promoters (“turbulators”) may be provided in the forward cavity. A portion of the air exits the forward cavity through an purge hole  86  extending through the inner band segment  52 . In the illustrated example, the purge hole  86  is sized such that about one-half of the mass flow entering the turbine vane  30  passes through the purge hole  86 , and is used to purge the rotor cavity of the turbine  10 . The remaining flow passes through a metering hole  88  and enters the impingement insert  60  through the inlet hole  84  in the inlet tube  82 . It exits the impingement insert  60  through impingement holes  80  as jets that cool the mid-cavity  56 . Next, the air exits the mid-cavity  56  through crossover holes  90  in the wall  92  between the mid-cavity  56  and the rear cavity  58 . It then cools the rear cavity  58  by convection. Although not shown, heat-transfer-enhancing structures such as fins, pins, turbulence promoters (“turbulators”) may be provided in the rear cavity  58 . The spent cooling air exits the rear cavity  58  through trailing edge passages  94 , such as the illustrated slots, or through film cooling holes (not shown) in the second stage vane  30 . 
     The turbine nozzle cooling configuration described herein is particularly useful when high pressure cooling air is available, but not in sufficient quantities to cool the entire nozzle with impingement. The design described herein combines turbulated lead edge cooling with mid-chord impingement and a warm bridge trailing edge to optimize the cooling and use of available air and pressure. The combination of these technologies enables a design that meets temperature goals which could not otherwise be met without higher cooling flow rates or extremely high turbulator enhancements. 
     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.