Abstract:
A turbine nozzle segment includes: (a) an arcuate outer band segment; (b) a hollow, airfoil-shaped turbine vane extending radially inward from the outer band segment; (c) a manifold cover secured to the outer band such that the manifold cover and the outer band segment cooperatively define an impingement cavity; and (d) an impingement blanket disposed in the impingement cavity, the impingement blanket having at least one impingement hole formed therethrough which is arranged to direct cooling air at the outer band segment. A method is provided for impingement cooling the outer band segment.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to gas turbine engine turbines and more particularly to methods for cooling turbine sections of such engines. 
         [0002]    A gas turbine engine includes a turbomachinery core having a high pressure compressor, combustor, and 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 a 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. 
         [0003]    The gas generator turbine includes annular arrays (“rows”) 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 minimized as much as possible. 
         [0004]    Prior art gas generator turbine nozzles have been cooled either using a “spoolie” fed manifold cover or a continuous impingement ring with a spoolie-fed airfoil insert. For the first system, air is fed into a manifold above the outer band, and then flows into the airfoil without directly cooling the outer band. The second configuration utilizes a separate impingement ring to cool the outer band, but this flow is susceptible to leakage through the gaps between adjacent nozzle segments. In either case, the turbine nozzle cooling is less efficient than desired. 
       BRIEF SUMMARY OF THE INVENTION 
       [0005]    These and other shortcomings of the prior art are addressed by the present invention, which provides independent impingement cooling for individual turbine nozzle outer band segments. 
         [0006]    According to one aspect of the invention, a turbine nozzle segment includes: (a) an arcuate outer band segment; (b) a hollow, airfoil-shaped turbine vane extending radially inward from the outer band segment; (c) a manifold cover secured to the outer band such that the manifold cover and the outer band segment cooperatively define an impingement cavity; and (d) an impingement blanket disposed in the impingement cavity, the impingement blanket having at least one impingement hole formed therethrough which is arranged to direct cooling air at the outer band segment. 
         [0007]    According to another aspect of the invention, a turbine nozzle assembly for a gas turbine engine includes: (a) a plurality of turbine nozzle segments arranged in an annular array, each turbine nozzle segment having: (i) an arcuate outer band segment; (ii) a hollow, airfoil-shaped turbine vane extending radially inwardly from the outer band segment; (iii) a manifold cover secured to the outer band such that the manifold cover and the outer band segment cooperatively define an impingement cavity; and (iv) an impingement blanket disposed in the impingement cavity, the impingement blanket having at least one impingement hole formed therethrough which is arranged to direct cooling air at the outer band segment; (b) an annular supporting structure surrounding the turbine nozzle segments; and (c) a plurality of generally cylindrical conduits, each conduit connecting one of the manifold covers in independent flow communication with the supporting structure. 
         [0008]    According to another aspect of the invention, a method is provided for cooling a turbine nozzle which includes an array of nozzle segments each having an arcuate outer band with a hollow, airfoil-shaped turbine vane extending radially inward therefrom. The method includes: (a) providing each of the outer bands with a closed impingement cavity having an impingement blanket disposed therein; (b) directing cooling air separately into the impingement cavities; (c) directing cooling air through one or more impingement holes in the impingement blanket against the outer band; and (d) exhausting the cooling air from the impingement cavity. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]    The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
           [0010]      FIG. 1  is a cross-sectional view of a high pressure turbine section of a gas turbine engine, constructed in accordance with an aspect of the present invention; 
           [0011]      FIG. 2  is a perspective view of a turbine nozzle shown in  FIG. 1 , with a manifold cover assembled thereto; 
           [0012]      FIG. 3  is perspective view of an impingement blanket; 
           [0013]      FIG. 4  is a perspective view of a manifold cover; and 
           [0014]      FIG. 5  is a perspective view of the impingement blanket of  FIG. 3  assembled to the manifold cover of  FIG. 4 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0015]    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. 
         [0016]    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. 
         [0017]    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 . 
         [0018]    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 . 
         [0019]    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 . 
         [0020]    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 . 
         [0021]    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. 
         [0022]    The second stage nozzle  28  is supported in part by mechanical connections to the first stage shroud hangers  46  and the shroud support  48 . Each second stage vane  30  is hollow so as to be able to receive cooling air in a known fashion. 
         [0023]      FIGS. 2-5  illustrate the construction of the second stage nozzle  28  in more detail.  FIG. 2  shows two individual nozzle segments  36  arranged side-by side, as they would be in the assembled gas generator turbine  10 . In the illustrated example, the nozzle segment  36  is a “singlet” casting which includes a segment  50  of the outer band  32 , a segment  52  of the inner band  34 , and a hollow second stage vane  30 . The radially outer end of each outer band segment  50  is closed by a manifold cover  54 . The manifold cover  54  (see  FIG. 4 ) is a unitary, slightly convex structure which has a lower peripheral edge  56  that matches the radially outer surface  58  of the outer band segment  50 , and includes an outwardly-extending inlet tube  60 . 
         [0024]    A plate-like impingement blanket  62 , best seen in  FIG. 3 , has a plurality of impingement holes  64  formed through it. It may be cast or fabricated from sheet metal. It is placed inside a recess  66  on the radially inner side of the manifold cover  54 , as seen in  FIG. 5 , and is secured thereto, for example by brazing, welding, fasteners, or adhesives. 
         [0025]    The manifold cover  54  is secured to the outer surface  58  of the outer band segment  50  so as to form an integral, sealed structure, with the sole inlet for air flow being the inlet tube  60 . As seen in  FIG. 1 , the manifold cover  54  and the outer band segment  50  cooperatively define an impingement cavity  68  which is divided into two sections by the impingement blanket  62 . 
         [0026]    When assembled, the inlet tube  60  is coupled to a generally cylindrical tube or conduit known as a “spoolie”  70 . The spoolie  70  penetrates the shroud support  48  to provide a pathway for cooling air into the interior of the second stage vanes  30 , as described in more detail below. One spoolie  70  is provided for each of the inlet tubes  60 . 
         [0027]    In operation, compressor discharge air (CDP), at the highest pressure in the compressor, or another suitable cooling air flow, is ducted to the shroud support  48  in a known manner. The CDP air enters the spoolies  66 , depicted by the arrows labeled “C” in  FIG. 1 . It then flows through the inlet tubes  60  into the individual impingement cavities  68  of each nozzle segment  36 . The cooling air exits the impingement holes  64  as a series of jets, depicted by the arrows “J”, which impinge against the outer band segment  50  and cool it. The spent impingement air is then exhausted to the interior of the turbine vane  30 , where is may be used to for additional cooling in a known manner. The area between the manifold cover  54  and the shroud support  48  is referred to as an outer band cavity  72 , and is purged by a separate air flow source. 
         [0028]    This configuration offers several advantages. By integrally joining the impingement blanket  62  to the manifold cover  54 , and by joining the manifold cover  54  to the outer band segment  50 , the outer band segment  50  can be impingement cooled using high pressure air without the associated inter-segment leakage penalties. This configuration then allows for the use of lower pressure air to purge the nozzle outer band cavities—as the air is at a lower pressure, the total amount of leakage flow will be reduced resulting in a lower performance penalty. 
         [0029]    The foregoing has described cooling arrangements for a turbine nozzle. 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.