Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
CA 02655689 2009-02-26
229199
TURBINE NOZZLE WITH INTEGRAL IMPINGEMENT BLANKET
BACKGROUND OF THE INVENTION
This invention relates generally to gas turbine engine turbines and more
particularly to methods for cooling turbine sections of such engines.
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.
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.
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.
-1-
CA 02655689 2009-02-26
229199
BRIEF SUMMARY OF THE INVENTION
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.
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.
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.
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
-2-
CA 02655689 2009-02-26
229199
holes in the impingement blanket against the outer band; and (d) exhausting
the cooling
air from the impingement 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:
Figure 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;
Figure 2 is a perspective view of a turbine nozzle shown in Figure 1, with a
manifold cover assembled thereto;
Figure 3 is perspective view of an impingement blanket;
Figure 4 is a perspective view of a manifold cover; and
Figure 5 is a perspective view of the impingement blanket of Figure 3
assembled
to the manifold cover of Figure 4.
DETAILED DESCRIPTION OF THE INVENTION
Referring to the drawings wherein identical reference numerals denote the same
elements throughout the various views, Figure 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.
-3-
CA 02655689 2009-02-26
229199
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 an 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 Figure 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.
-4-
CA 02655689 2009-02-26
229199
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. Each second stage
vane 30 is
hollow so as to be able to receive cooling air in a known fashion.
Figures 2-5 illustrate the construction of the second stage nozzle 28 in more
detail. Figure 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 Figure 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.
A plate-like impingement blanket 62, best seen in Figure 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
Figure 5, and is secured thereto, for example by brazing, welding, fasteners,
or adhesives.
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 Figure 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.
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.
-5-
CA 02655689 2009-02-26
229199
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
Figure 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
ofjets, 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.
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.
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.
-6-
