U.S. patent application number 12/892056 was filed with the patent office on 2012-03-29 for conduction pedestals for a gas turbine engine airfoil.
Invention is credited to Amanda Jean Leamed, Brandon W. Spangler.
Application Number | 20120076660 12/892056 |
Document ID | / |
Family ID | 44719513 |
Filed Date | 2012-03-29 |
United States Patent
Application |
20120076660 |
Kind Code |
A1 |
Spangler; Brandon W. ; et
al. |
March 29, 2012 |
CONDUCTION PEDESTALS FOR A GAS TURBINE ENGINE AIRFOIL
Abstract
An airfoil for a gas turbine engine includes an airfoil which
defines a leading edge cavity and a forward cavity between a
pressure side wall and a suction side wall, the leading edge cavity
at least partially defined by a leading edge wall which extends
between the pressure side wall and the suction side wall. A rib
between the pressure side wall and the suction side wall separates
the forward cavity and the leading edge cavity. A pedestal extends
between the leading edge wall and the rib.
Inventors: |
Spangler; Brandon W.;
(Vernon, CT) ; Leamed; Amanda Jean; (Manchester,
CT) |
Family ID: |
44719513 |
Appl. No.: |
12/892056 |
Filed: |
September 28, 2010 |
Current U.S.
Class: |
416/223R |
Current CPC
Class: |
F05D 2260/221 20130101;
F05D 2260/201 20130101; F05D 2240/121 20130101; F05D 2260/202
20130101; F05D 2260/2214 20130101; F01D 5/187 20130101 |
Class at
Publication: |
416/223.R |
International
Class: |
F04D 29/38 20060101
F04D029/38 |
Claims
1. An airfoil for a gas turbine engine comprising: a pressure side
wall and a suction side wall which define a leading edge cavity and
a forward cavity between said pressure side wall and said suction
side wall, said leading edge cavity at least partially defined by a
leading edge wall which extends between said pressure side wall and
said suction side wall; a rib between said pressure side wall and
said suction side wall to at least partially divide said forward
cavity and said leading edge cavity; and a pedestal which extends
between said leading edge wall and said rib.
2. The airfoil as recited in claim 1, wherein said pedestal is
aligned along an axis which extends toward a high temperature area
in a stagnation region of said leading edge wall.
3. The airfoil as recited in claim 2, further comprising a second
pedestal aligned along a second axis different than said axis.
4. The airfoil as recited in claim 1, wherein said rib at least
partially defines an impingement leading edge.
5. The airfoil as recited in claim 4, wherein said rib defines a
multiple of cooling holes which communicate a cooling flow from
said forward cavity into said leading edge cavity through said rib
then through a multiple of leading edge cooling holes through said
leading edge.
6. The airfoil as recited in claim 1, wherein said rib at least
partially defines a radial flow leading edge.
7. The airfoil as recited in claim 6, wherein said leading edge
defines a multiple of cooling holes which communicate a cooling
flow from within said leading edge cavity through a multiple of
leading edge cooling holes through said leading edge.
8. The airfoil as recited in claim 1, wherein said airfoil at least
partially defines a turbine vane.
9. The airfoil as recited in claim 1, wherein said airfoil at least
partially defines a turbine blade.
10. An airfoil for a gas turbine engine comprising: a pressure side
wall and a suction side wall which defines a leading edge cavity
and a forward cavity between said pressure side wall and said
suction side wall, said leading edge cavity at least partially
defined by a leading edge wall which extends between said pressure
side wall and said suction side wall; a rib between said pressure
side wall and said suction side wall to at least partially divide
said forward cavity and said leading edge cavity; and a multiple of
pedestals which extend between said leading edge wall and said rib,
said multiple of pedestals arrayed along a length of said airfoil
between a first end portion and a second end portion.
11. The airfoil as recited in claim 10, wherein each of said
multiple of pedestals are aligned along an axis which extends
toward a high temperature area in a stagnation region of said
leading edge wall.
12. The airfoil as recited in claim 10, wherein a first set of said
multiple of pedestals are aligned along a first axis which extends
toward a first high temperature area in a stagnation region of said
leading edge and a second set of said multiple of pedestals are
aligned along a second axis which extends toward a second high
temperature area in the stagnation region of said leading edge.
13. The airfoil as recited in claim 10, wherein each of said
multiple of pedestals are transverse to said rib.
14. The airfoil as recited in claim 10, wherein said airfoil at
least partially defines a turbine vane.
15. The airfoil as recited in claim 10, wherein said airfoil at
least partially defines a turbine blade.
Description
BACKGROUND
[0001] The present disclosure relates to a gas turbine engine, and
more particularly to an airfoil cooling arrangement.
[0002] A gas turbine engine includes a compressor section that
compresses air then channels the compressed air to a combustor
section wherein the compressed airflow is mixed with fuel and
ignited to generate high temperature combustion gases. The
combustion core gases flow downstream through a turbine section
which extracts energy therefrom to power the compressor section and
a fan section. Since the combustion core gases are at a high
temperature, turbine vanes and turbine blades within the turbine
section may have relatively high heat loads at the leading
edges.
SUMMARY
[0003] An airfoil for a gas turbine engine according to an
exemplary aspect of the present disclosure includes a pressure side
wall and a suction side wall which define a leading edge cavity and
a forward cavity between the pressure side wall and the suction
side wall, with the leading edge cavity at least partially defined
by a leading edge wall which extends between the pressure side wall
and the suction side wall. A rib between the pressure side wall and
the suction side wall separates the forward cavity and the leading
edge cavity. A pedestal extends between the leading edge wall and
the rib.
[0004] An airfoil for a gas turbine engine according to an
exemplary aspect of the present disclosure includes a multiple of
pedestals which extend between a leading edge and a rib, the
multiple of pedestals arrayed along a length of the airfoil between
a first end portion and a second end portion.
BRIEF DESCRIPTION OF THE DRAWINGS
[0005] Various features will become apparent to those skilled in
the art from the following detailed description of the disclosed
non-limiting embodiments. The drawings that accompany the detailed
description can be briefly described as follows:
[0006] FIG. 1 is a general schematic partial fragmentary view of an
exemplary gas turbine engine embodiment for use with the present
invention;
[0007] FIG. 2 is a perspective view of a vane;
[0008] FIG. 3 is a sectional view of an airfoil;
[0009] FIG. 4 is a perspective partial fragmentary view of an
airfoil with an impingement flow leading edge;
[0010] FIG. 5 is a perspective partial fragmentary view of an
airfoil with a radial flow leading edge;
[0011] FIG. 6 is a sectional view of a leading edge of an airfoil
with a pedestal according to one non-limiting embodiment;
[0012] FIG. 7 is a sectional view of a RELATED ART airfoil leading
edge which illustrates a temperature gradient therein to determine
an associated conduction path axis;
[0013] FIG. 8 is a sectional view of a RELATED ART airfoil leading
edge which illustrates a temperature gradient therein to locate the
pedestals of FIG. 7;
[0014] FIG. 9 is a sectional view of the airfoil leading edge of
FIG. 6 which illustrates a temperature gradient therein as reduced
due to the pedestals;
[0015] FIG. 10 is a sectional view of a leading edge of an airfoil
with pedestals according to one non-limiting embodiment; and
[0016] FIG. 11 is a sectional view of a RELATED ART airfoil leading
edge which illustrates a temperature gradient therein to determine
associated conduction path axes to locate the pedestals of FIG.
10.
DETAILED DESCRIPTION
[0017] FIG. 1 schematically illustrates a gas turbine engine 10
which generally includes a fan section 12, a compressor section 14,
a combustor section 16, and a turbine section 18. Within and aft of
the combustor section 16, engine components are typically cooled
due to intense temperature of the combustion core gases. While a
two spool high bypass turbofan engine is schematically illustrated
in the disclosed non-limiting embodiment, it should be understood
that the disclosure is applicable to other gas turbine engine
configurations.
[0018] At least some stages of the turbine rotor blades 22 and
turbine stator vanes 24 within the turbine section 18, for example,
may be cooled with a cooling airflow typically sourced with a bleed
airflow from the compressor section 14 at temperature lower than
the core gas within the turbine section 18. The cooling airflow
passes through at least one cooling circuit flow path 26 (FIG. 2)
to transfer thermal energy from the component to the cooling
airflow.
[0019] Each cooling circuit flow path 26 may be disposed in any
component that requires cooling, and in most cases the component
receives cooling airflow therethrough as the external surface
thereof is exposed to combustion core gases. In the illustrated
embodiment and for purposes of giving a detailed example, the
cooling circuit flow path 26 will be described herein as being
disposed within a portion of an airfoil 32 such as that of a stator
vane 24 or rotor blade 22. It should be understood, however, that
the cooling circuit flow path 26 is not limited to these
applications and may be utilized within other areas such as liners,
seals, and other structures with stagnation regions exposed to high
temperature core gas flow.
[0020] With reference to FIG. 2, the cooling circuit flow path 26
communicates with a multiple of cavities, for example 34A-34B shown
in FIG. 3, formed within the airfoil 32. The multiple of cavities
34A-34B direct cooling airflow which may include air received from
the compressor section into high temperature areas of the airfoil
32.
[0021] The airfoil 32 is defined by an outer airfoil wall surface
40 between a leading edge 36 and a trailing edge 42. The outer
airfoil wall surface 40 typically has a generally concave shaped
portion forming a pressure side 40P and a generally convex shaped
portion forming a suction side 40S which are connected by a leading
edge wall 40L at the leading edge 36. The outer airfoil wall
surface 40 is longitudinally defined to span a first end portion 46
and a second end portion 48. The end portions 46, 48 may include
features to mount the airfoil to other structures such as engine
static structure or rotor disk. For example, the end portions 46,
48 for a vane may include outer vane platforms and for a blade may
include an attachment section and a blade tip. It should be
understood that various component arrangement may likewise be
utilized with the present invention.
[0022] With reference to FIG. 3, the forward cavity 34A is
generally defined by a first rib 54 just aft of the leading edge
36. The first rib 54 separates the forward cavity 34A from a
leading edge cavity 56 defined at least partially by the outer
airfoil wall surface 40 and often referred to as a "peanut" cavity.
The first rib 54 may, for example, at least partially define an
impingement leading edge 62 (FIG. 4) or a radial flow leading edge
64 (FIG. 5) which may span a portion of or the entire length of the
airfoil 32. That is, the pedestals 60 may be specifically located
along the entire airfoil 32 span or a select portion or portions
thereof.
[0023] The leading edge cavity 56 includes the multiple of
pedestals 60 which are transverse to and extend between the leading
edge 36 and the first rib 54. It should be understood that any
number of pedestals 60 may be so positioned. The pedestals 60
provide an additional thermal conductive path along a conduction
path axis H (FIG. 6) from the leading edge 36 to the first rib 54
to reduce the temperature of the leading edge 36 as the leading
edge 36 may otherwise be hundreds of degrees hotter than the
pressure side 40P and suction side 40S of the airfoil 32 due to
higher external heat transfer coefficients at the stagnation region
S (FIG. 7). It should be understood that the stagnation region S is
a region within which the combustion gas flow Mach number may be
relatively low such that a temperature concentration occurs.
[0024] For the impingement leading edge 62 cooling scheme (FIG. 4)
the first rib 54 may define a multiple of cooling holes 66 which
communicate a cooling flow from the forward cavity 34A into the
leading edge cavity 56 through the first rib 54 then out through a
multiple of leading edge cooling holes 68. That is, the cooling
flow is communicated generally along the pedestals 60. For the
radial flow leading edge 64 cooling scheme (FIG. 5) the cooling
flow from within the leading edge cavity 56 passes transverse to
the pedestals 60 and out through a multiple of leading edge cooling
holes 70. It should be understood that various such cooling schemes
will benefit from the pedestals 60.
[0025] The pedestals 60 reduce leading edge 36 temperatures mainly
from the enhanced conduction effects of the pedestals 60 from the
leading edge 36 to the first rib 54 (FIGS. 8 and 9). In addition,
for radial flow leading edges (FIG. 5), a portion of the metal
temperature reduction is achieved by the enhancement of the
internal heat transfer coefficient as coolant flow passes over the
pedestals 60. The lower temperature at the stagnation region
beneficially results in, for example, a higher oxidation, local
creep, and Thermal Mechanical Fatigue (TMF) capability.
[0026] The pedestals 60 may be selectively oriented at a multiple
of different angles in the leading edge cavity 56 to achieve the
desired thermal reduction effect. That is, the pedestals 60-1, 60-2
may be aligned along conduction path axes H1, H2 (FIG. 10) which
extend into the highest temperature areas in the stagnation region
of the leading edge 36 (FIG. 11) to facilitate a more direct heat
transfer from the leading edge 36 to the first rib 54. It should be
understood that the axes H1, H2 may change along the span of the
airfoil 32. The relative positions of the pedestals 60-1, 60-2 may
thereby also change along the span to correspond therewith.
[0027] The manufacture of the pedestals 60 may be achieved by a
proprietary Fugitive Core Process which uses thermoplastic inserts
to create a one piece core with multiple pull angles as developed
by Alcoa Howmet of Cleveland Ohio USA. Generally, sacrificial
thermoplastic pieces make up the rib and leading edge pedestals;
the thermoplastic pieces are assembled into the core die and core
material is injected around the thermoplastic pieces; the
thermoplastic pieces are melted, leaving voids in finished core;
and metal fill voids in core to form pedestals in the finished
part.
[0028] It should be understood that like reference numerals
identify corresponding or similar elements throughout the several
drawings. It should also be understood that although a particular
component arrangement is disclosed in the illustrated embodiment,
other arrangements will benefit herefrom.
[0029] Although particular step sequences are shown, described, and
claimed, it should be understood that steps may be performed in any
order, separated or combined unless otherwise indicated and will
still benefit from the present disclosure.
[0030] The foregoing description is exemplary rather than defined
by the limitations within. Various non-limiting embodiments are
disclosed herein, however, one of ordinary skill in the art would
recognize that various modifications and variations in light of the
above teachings will fall within the scope of the appended claims.
It is therefore to be understood that within the scope of the
appended claims, the disclosure may be practiced other than as
specifically described. For that reason the appended claims should
be studied to determine true scope and content.
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