JP2009209936A - Turbine nozzle with integral impingement blanket - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a turbine nozzle with an integral impingement blanket. <P>SOLUTION: A turbine nozzle segment has (a) an arched outer band segment 50, (b) a hollow blade-shaped turbine blade 30 extending inward in the radial direction from the outer band segment 50, (c) a manifold cover 54 fixed to the outer band segment 50 such that the manifold cover 54 and the outer band segment 50 cooperatively define an impingement cavity, and (d) an impingement blanket arranged in the impingement cavity for penetratingly forming at least one impingement hole to be constituted so as to direct cooling air at the outer band segment 50. <P>COPYRIGHT: (C)2009,JPO&INPIT

Description

  The present invention relates generally to gas turbine engine turbines, and more particularly to a method for cooling the turbine portion of such engines.

  The gas turbine engine includes a turbomachine core having a high pressure compressor, a combustor, and a high pressure or gas generator in a DC relationship. The core is operable in a known manner and generates a temporary gas stream. In a turbojet or turbofan engine, core exhaust is guided through the nozzles 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 an annular array ("row") of stationary vanes or nozzles that guide the gas exiting the combustor into rotor blades or buckets. A row of nozzles and a row of blades collectively form a “stage”. Generally, two or more stages are used in a direct current relationship. These components operate in very high temperature environments and need to be cooled by an air flow to ensure proper service life. In general, the air used for cooling is extracted at one or more points in the compressor. These bleed airflows represent the net work and / or thrust loss of the thermodynamic cycle. These increase the fuel consumption rate (SFC) and should usually be minimized as much as possible.

  Prior art gas generator turbine nozzles have been cooled using either a “spoolie” feed manifold cover or a continuous impingement ring with a spur feed airfoil insert. For the first scheme, air is supplied to the manifold above the outer band and 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 prone to leak through gaps between adjacent nozzle segments. In either case, the efficiency of turbine nozzle cooling is worse than desired.

  These and other shortcomings of the prior art are overcome by the present invention which provides independent impingement cooling of individual turbine nozzle outer band segments.

  In accordance with one aspect of the present invention, a turbine nozzle segment includes: (a) an arcuate outer band segment; (b) a hollow airfoil shaped turbine blade extending radially inward from the outer band segment; A manifold cover secured to the outer band segment, wherein the manifold cover and the outer band segment jointly define a collision cavity; and (d) a collision blanket disposed within the collision cavity. And a collision blanket formed with at least one collision hole formed therethrough configured to guide cooling air to the outer band segment.

  In accordance with another aspect of the present invention, a turbine nozzle assembly of a gas turbine engine comprises (a) a plurality of turbine nozzle segments arranged in an annular array, each turbine nozzle segment comprising: (i) an arcuate outer side. A band segment; (ii) a hollow airfoil shaped turbine blade extending radially inward from the outer band segment; and (iii) a manifold cover secured to the outer band segment, the manifold cover and the A manifold cover with outer band segments jointly defining a collision cavity; and (iv) a collision blanket disposed in the collision cavity, the at least one configured to guide cooling air to the outer band segment The turbine nozzle having the collision blanket having one collision hole formed therethrough And segment, and a plurality of substantially cylindrical conduit that bind (b) the turbine and an annular supporting structure surrounding the nozzle segments, (c) each said supporting one of said manifold cover independently communicating state structure.

  In accordance with another aspect of the present invention, a method for cooling a turbine nozzle is provided that includes an array of nozzle segments each having an arcuate outer band segment with a hollow airfoil shaped turbine blade extending radially inward. Is done. The method includes: (a) providing a closed collision cavity with an impact blanket disposed therein for each of the outer band segments; (b) separately guiding cooling air to the collision cavity; c) guiding cooling air to the outer band segment through one or more impingement holes of the impingement blanket; and (d) discharging cooling air from the impingement cavity.

The invention can best be understood by reference to the following description taken in conjunction with the drawings in the accompanying drawings.
1 is a cross-sectional view of a high pressure turbine portion of a gas turbine engine configured in accordance with an aspect of the present invention. FIG. 2 is a perspective view of the turbine nozzle shown in FIG. 1 with a manifold cover attached. It is a perspective view of a collision blanket. It is a perspective view of a manifold cover. FIG. 5 is a perspective view of the collision blanket of FIG. 3 attached to the manifold cover of FIG. 4.

  Referring to the drawings in which like reference numbers indicate like elements throughout the Figures, FIG. 1 illustrates a portion of a gas generator turbine 10 that is part of a known type of gas turbine engine. The function of the gas generator turbine 10 is to extract energy from the hot pressurized combustion gas of an upstream combustor (not shown) and to convert that energy into 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 the work turbine is positioned downstream of the gas generator turbine 10 and connected to the output shaft. However, the principles described herein are equally applicable to turbine engines used in other vehicle or stationary applications as well as turboprop, turbojet, and turbofan engines.

  The gas generator turbine 10 includes a plurality of circumferentially spaced supports supported between an arcuately segmented first stage outer band 16 and an arcuately segmented first stage inner band 18. The first stage nozzle 12 including the first stage blade 14 of the hollow airfoil shape is included. The first stage blades 14, the first stage outer band 16, and the first stage inner band 18 are arranged in a plurality of circumferentially adjacent nozzle segments that collectively form a complete 360 ° assembly. First stage outer band 16 and first stage inner band 18 define the outer and inner radial flow path boundaries, respectively, for the hot gas flow through first stage nozzle 12. The first stage blade 14 is configured to optimally guide the combustion gas to the first stage rotor 20.

  The first stage rotor 20 includes an array of airfoil shaped first stage turbine blades 22 extending outwardly from a first stage disk 24 that rotates about an engine centerline axis. The segmented arcuate first stage shroud 26 closely surrounds the first stage turbine blades 22 so as to define an outer radial flow path boundary for hot gas flow through the first stage rotor 20. Composed.

  The second stage nozzle 28 is positioned downstream of the first stage rotor 20 and is supported between an arched second stage outer band 32 and an arched second stage inner band 34. A plurality of hollow airfoil-shaped second stage blades 30 that are spaced apart in the circumferential direction. The second stage blade 30, the second stage outer band 32, and the first stage inner band 34 are arranged in a plurality of circumferentially adjacent nozzle segments 36 (see FIG. 2) that collectively form a complete 360 ° assembly. Is done. Second stage outer band 32 and second stage inner band 34 define outer and inner radial flow path boundaries for the hot gas flow flowing through second stage turbine nozzle 28, respectively. The second stage blade 30 is configured to optimally guide the combustion gas to the second stage rotor 38.

  The second stage rotor 38 includes a radial array of airfoil shaped second stage turbine blades 40 extending radially outward from a second stage disk 42 that rotates about an engine centerline axis. The segmented arcuate second stage shroud 44 closely surrounds the second stage turbine blade 40 to define an outer radial flow path boundary for the hot gas flow through the second stage rotor 38. Composed.

  The segments of the first stage shroud 26 are, for example, of the arcuate first stage shroud hanger 46 that is also carried by the arcuate shroud support 48 using the illustrated hooks, rails, and C-shaped clips in a known manner. Supported by an array.

  The second stage nozzle 28 is partially supported by a mechanical connection with the first stage shroud hanger 46 and the shroud support 48. Each second stage blade 30 is hollow so as to receive cooling air in a known manner.

  2 to 5 show the structure of the second stage nozzle 28 in more detail. FIG. 2 shows two individual nozzle segments 36 arranged side by side for entry into the assembled gas generator turbine 10. In the illustrated example, nozzle segment 36 is a “singlet” casting that includes segment 50 of outer band 32, segment 52 of inner band 34, and hollow second stage vane 30. The radially outer end of each outer band segment 50 is sealed with a manifold cover 54. The manifold cover 54 (see FIG. 4) is an integral, slightly convex structure having a lower peripheral edge 56 that conforms to the radially outer surface 58 of the outer band segment 50 and includes an outwardly extending inflow tube 60. Prepare.

  The plate-shaped collision blanket 62 most often seen in FIG. 3 has a plurality of collision holes 64, which are formed through the collision blanket 62. This blanket can be cast or manufactured from sheet metal. This is located within a recess 66 radially inward of the manifold cover 54, as seen in FIG. 5, and secured by, for example, brazing, welding, fasteners, or adhesive.

  The manifold cover 54 is secured to the outer surface 58 of the outer band segment 50 so as to form an integral hermetic structure with only one inlet for airflow, which is the inlet tube 60. As seen in FIG. 1, the manifold cover 54 and the outer band segment 50 jointly define a collision cavity 68 that is divided into two parts by a collision blanket 62.

  During assembly, the inflow tube 60 is connected to a generally cylindrical tube or conduit known as a “spulley” 70. The spur 70 penetrates the shroud support 48 and forms a path for cooling air inside the second stage blade 30 as will be described in detail later. One spur 70 is provided for each inflow pipe 60.

  In operation, compressor exhaust at the highest compressor pressure (CDP: compressor discharge pressure), or another suitable cooling air flow, is delivered to the shroud support 48 in a known manner. CDP air enters the pulley 70 as indicated by arrow “C” in FIG. Air then flows through the inlet pipe 60 into the individual impingement cavities 68 of each nozzle segment 36. The cooling air exits the impingement hole 64 as a series of jets as shown by arrow “J” and impinges on the outer band segment 50 to cool it. Thereafter, the used impinging air is exhausted into the turbine blade 30, at which time the impinging air can be used for further cooling in a known manner. The area between the manifold cover 54 and the shroud support 48 is called the outer band cavity 72 and is purged by a separate air flow source.

  This configuration offers several advantages. By joining the impact blanket 62 integrally with the manifold cover 54 and further joining the manifold cover 54 with the outer band segment 50, the outer band can be used with high pressure air without the disadvantage of leakage between segments. The segment 50 can be impact cooled. And this structure allows low pressure air to be used to purge the nozzle outer band cavity, and the air is at a low pressure, reducing the total amount of leakage flow and consequently reducing performance disadvantages. It will be.

  The method for cooling the turbine nozzle has been described above. Although the invention has been described herein with reference to specific embodiments, it will be apparent to those skilled in the art that various modifications can be made without departing from the scope of the appended claims. Is possible. Accordingly, the present invention is defined by the appended claims, and the description of the preferred embodiment of the invention and the best mode for carrying out the invention is not intended to be limiting, but merely for illustrative purposes. It was made.

Claims (8)

A turbine nozzle segment (36) comprising:
(A) an arched outer band segment (50);
(B) a hollow airfoil shaped turbine blade (30) extending radially inward from the outer band segment (50);
(C) a manifold cover (54) secured to the outer band segment (50), wherein the manifold cover (54) and the outer band segment (50) jointly define a collision cavity (68); A manifold cover (54);
(D) at least one impact hole (64) configured to guide cooling air to the outer band segment (50), the impact blanket (62) being disposed within the impact cavity (68); A turbine nozzle segment (36) consisting of said impingement blanket (62) formed therethrough.   The turbine nozzle segment (36) of any preceding claim, having a plurality of impact holes (64) in which the impact blanket (62) is formed. (A) having a radially inwardly facing recess (66) in which the manifold cover (54) is formed;
A turbine nozzle segment (36) according to claim 1 or claim 2, wherein the collision blanket (62) is a plate secured to the manifold cover (54) to close the recess (66).   The turbine nozzle segment (36) of claim 3, wherein the impingement blanket (62) is brazed to the manifold cover (54).   The turbine nozzle segment (36) according to any of the preceding claims, wherein the manifold cover (54) is brazed to the outer band segment (50).   The turbine nozzle segment (36) of claim 1, wherein the manifold cover (54) includes an inflow tube (60) extending radially outward.   The turbine nozzle segment (36) according to any of the preceding claims, further comprising an arcuate inner band segment (52) disposed at a radially inner end of the turbine blade (30). A method of cooling a turbine nozzle (28) comprising an array of nozzle segments each having an arcuate outer band segment (50) with a hollow airfoil shaped turbine blade (30) extending radially inwardly. ,
(A) providing a closed collision cavity (68) with an impact blanket (62) disposed therein to each of said outer band segments (50);
(B) separately guiding cooling air to the collision cavity (68);
(C) guiding cooling air to the outer band segment (50) through one or more collision holes (64) of the collision blanket (62);
(D) discharging the cooling air from the collision cavity (68).