U.S. patent application number 11/787556 was filed with the patent office on 2009-03-19 for turbine blade with cooling breakout passages.
This patent application is currently assigned to Florida Turbine Technologies, Inc.. Invention is credited to Joseph Brostmeyer.
Application Number | 20090074576 11/787556 |
Document ID | / |
Family ID | 40454663 |
Filed Date | 2009-03-19 |
United States Patent
Application |
20090074576 |
Kind Code |
A1 |
Brostmeyer; Joseph |
March 19, 2009 |
Turbine blade with cooling breakout passages
Abstract
A turbine airfoil with a plurality of breakout passages located
just beneath a thermal barrier coating or just beneath the metal
surface of the airfoil. The breakout passages are connected to an
internal cooling air passage and allow for film cooling air to flow
when a surface over the breakout passage has been chipped or eroded
away to provide for additional cooling to the damaged airfoil
surface. The breakout passages are formed during the casting
process of the airfoil using a plurality of molding pieces. The
breakout passages are formed along a direction of the pulling
direction of the respective mold piece. This allows for the
solidified airfoil to be easily removed from the mold assembly.
Inventors: |
Brostmeyer; Joseph;
(Jupiter, FL) |
Correspondence
Address: |
Joseph Brostmeyer;Florida Turbine Technologies, Inc.
Suite 110, 100 Marquette Road
Jupiter
FL
33458
US
|
Assignee: |
Florida Turbine Technologies,
Inc.
Jupiter
FL
|
Family ID: |
40454663 |
Appl. No.: |
11/787556 |
Filed: |
April 17, 2007 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
60794173 |
Apr 20, 2006 |
|
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|
Current U.S.
Class: |
416/95 |
Current CPC
Class: |
F01D 5/186 20130101;
F01D 5/187 20130101 |
Class at
Publication: |
416/95 |
International
Class: |
F01D 5/08 20060101
F01D005/08 |
Claims
1. A turbine airfoil for use in a gas turbine engine, the airfoil
comprising: An airfoil surface for exposure to a hot gas flow
through the turbine; An internal cooling air supply passage to
channel cooling air through the airfoil and provide cooling; A
plurality of cooling breakout passages connected to the cooling air
supply passage on an upstream end and extending toward the airfoil
surface on the downstream end, the cooling breakout passages being
closed to cooling air flow when the airfoil surface is intact and
open to cooling air flow when the airfoil surface has eroded away;
and, The turbine airfoil being formed from a molding process that
uses a plurality of mold pieces and the plurality of breakout
passages are aligned with a pulling direction of the mold piece in
which the breakout passages are formed.
2. The turbine airfoil of claim 1, and further comprising: The
airfoil surface above the breakout passages includes a thermal
barrier coating; and, The breakout passages extend through the
airfoil wall and to the thermal barrier coating such that when the
thermal barrier coating spalls the breakout passage below the
spallation will open to cooling air flow.
3. The turbine airfoil of claim 1, and further comprising: The
breakout passages extend to a point just under the airfoil metal
surface such that erosion of the metal airfoil surface just above
the breakout passage must occur before the breakout passage is open
to cooling air flow.
4. The turbine airfoil of claim 3, and further comprising: A
thermal barrier coating is applied to the airfoil surface over the
breakout passages.
5. A process of casting a turbine airfoil having an internal
cooling air supply passage and a plurality of cooling breakout
passages, the process comprising the steps of: Casting the airfoil
using a multiple piece mold assembly in which each mold piece has a
pulling direction; and, Casting the breakout passages along a
direction substantially parallel to the pulling direction of the
respective mold piece in which the passages are cast.
6. The process of casting a turbine airfoil of claim 5, and further
comprising the step of: Casting the airfoil in a mold assembly in
which the mold pieces are joined together along a line forming a
midpoint of the cross section of the airfoil.
7. The process of casting a turbine airfoil of claim 5, and further
comprising the step of: Casting the breakout passages so that the
passages end at a point just below the metal airfoil surface.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims the benefit to an earlier filed
Provisional Patent Application 60/794,173 filed on Apr. 20, 2006
and entitled TURBINE BLADE WITH COOLING BREAKOUT PASSAGES.
BACKGROUND OF THE INVENTION
[0002] 1. Field of the Invention
[0003] The present invention relates to gas turbine airfoils, and
more specifically to turbine airfoils having a thermal barrier
coating and internal cooling passages.
[0004] 2. Description of the Related Art Including Information
Disclosed Under 37 CFR 1.97 and 1.98
[0005] In gas turbine engines, airfoils (both moving blades and
stationary vanes) include cooling fluid passages within the airfoil
that form a closed (or open) cooling passage. A thermal barrier
coating (TBC) can also be applied to an outer surface of the
airfoil to provide a heat shield and prevent damage to the airfoil
due to high temperatures.
[0006] Spallation of the TBC is a very common in gas turbine
engines. When the TBC spalls, a small portion of the coating is
broken off from the airfoil, exposing the substrate metal or
airfoil surface below the TBC to extremely high gas temperatures of
the gas turbine. Usually, the high gas temperature is higher than
the melting temperature of the airfoil, especially in the first and
second stages of the turbine. Thus, when spallation--or other means
such as erosion or oxidation or foreign object damage--removes a
piece of the TBC, the metal substrate is exposed to the high gas
temperature and will melt away with time.
[0007] U.S. Pat. No. 6,039,537 issued to Scheurlen on Mar. 21, 2000
shows a gas turbine blade with internal cooling passages and a
thermal barrier coating (TBC) applied to the outer surface. Smaller
cooling air passages are located between the internal cooling
passages and the TBC. Some of the smaller cooling air passages are
covered up by the TBC such that the passage is closed to cooling
air flow. When the TBC above the opening of a smaller cooling air
passage is broken away (or, eroded or oxidized), the passage
becomes open and cooling air flows through the passage from the
internal cooling air passage out onto the surface of the airfoil
around the removed section of substrate, allowing for additional
cooling of the airfoil. As disclosed in the Scheurlen patent, "in
the event of a failure of the heat insulating layer system in the
effected region of the turbine blade, provision is made for
additional cooling by virtue of the fact that the heat insulating
system which breaks off opens the closed bore and enables a
coolant, which is operationally admitted to the interior space
anyway, to flow through the opened bore and thus intensify the
cooling of the affected region. The heat insulating layer system is
constructed in such a way that the use of the closed bore for
cooling the turbine blade is not necessary in the case of an
undamaged heat insulating layer system. The demand for coolant can
therefore be adapted to the protective properties of the heat
insulating layer system and be kept at a correspondingly low level.
In addition, the provision of corresponding bores to be closed by
the heat insulating layer system enables the turbine blade to be
reliably cooled by repeated discharge of coolant from the interior
space, and thus protected against undesirable failure even in the
event of a loss of the heat insulating layer system", and "all of
the bores are disposed in the substrate in such a way that the
substrate is uniformly cooled when the hot gas flow flows around
it, if the heat insulating layer system is open previously closed
bores when a cooling fluid drawn off through the bores into the gas
flow is fed to the interior space", and "such a structure also
permits monitoring of the turbine blade with regard to the
integrity of the heat insulating layer system by the inflow of the
coolant being measured and compared with a value which must appear
when the heat insulating layer system is intact, with all
corresponding bores being closed".
[0008] U.S. Pat. No. 5,269,653 issued to Evans on Dec. 14, 1993 and
entitled AEROFOIL COOLING discloses a turbine airfoil with a
leading edge having a row of spaced blank passages (# 30 is the
Evans patent) having blank ends that end just below the leading
edge airfoil surface. When a piece of the TBC erodes away, the hot
gas will erode away the blank ends of a passage and open the
passage so that cooling air will flow through the passage and out
onto the leading edge surface.
[0009] U.S. Pat. No. 6,749,396 B2 issued to Barry et al on Jun. 15,
2004 and entitled FAILSAFE FILM COOLED WALL discloses a gas turbine
engine with cooling where a number of failsafe film cooling holes
are located below a TBC in areas of high risk of thermal barrier
coating spallation and that do not permit cooling flow through the
holes unless the thermal barrier coating has eroded by spallation
for opening the outlet ends of the holes during normal operation of
the engine.
[0010] An object of the present invention is to provide for an air
cooled airfoil with small film cooling holes that are normally
closed to air flow but are opened when the portion of the airfoil
around the hole becomes too hot.
[0011] Another object of the present invention is to provide for an
air cooled airfoil with film cooling holes aligned with a pulling
direction of the mold.
SUMMARY OF THE INVENTION
[0012] The present invention is an air cooled turbine blade for a
gas turbine engine, the turbine blade having a cooling air passage
therethrough for channeling cooling air through the blade, and a
thermal barrier coating (TBC) or oxidation coating (or no coating
at all) applied to the exterior of the blade to protect the metal
substrate of the blade form damage due to a high temperature of the
gas. Located between the internal cooling air passage and the TBC
are small cooling air passages that form a closed cooling air path
and extend from the internal cooling air passage to a point below
the surface of the metal substrate on which the TBC is applied,
forming a closed cooling air channel in the blade metal
substrate.
[0013] When a piece of the TBC is broken away from the airfoil, the
metal substrate below the TBC is then exposed to the high gas
temperature. The high gas temperature then begins to melt away the
metal substrate around the exposed TBC-less area. Eventually, the
metal substrate melts to the point where the resulting hole joins
the small cooling air passage just below the substrate such that
cooling air flowing through the internal passage is also allowed to
flow through the smaller cooling air passage and out onto the
exposed surface of the metal substrate. The original closed cooling
air passage now becomes an open cooling air passage as cooling air
from inside the blade is directed out of the blade through the
small cooling air passage to cool the newly exposed area of the
metal substrate. Thus, further damage to the metal substrate of the
airfoil due to the missing TBC is prevented while allowing the
turbine engine to continue under full operating load until a later
time when the damage can be discovered and the TBC repaired.
[0014] The turbine airfoil having the breakout passages is formed
from a casting process with a mold formed from multiple pieces.
Each mold piece has a distinct pulling direction in which the mold
pieces are pulled away after the airfoil has been cast. The
breakout passages are formed during the casting process. Each mold
piece forms the breakout passages along a direction parallel to the
pulling direction of the mold piece. This allows for the easy
removal of the solidified airfoil after the molding process.
BRIEF DESCRIPTION OF THE DRAWINGS
[0015] FIG. 1 shows a turbine blade with an internal closed cooling
air passage and smaller cooling air passages extending from the
internal passage to a point near the surface of the metal substrate
and TBC.
[0016] FIG. 2 shows a cross sectional view of the airfoil with a
metal substrate and an undamaged TBC on the exterior surface of the
metal, and several smaller cooling air passages extending from the
internal cooling passage towards the outer surface of the metal
substrate.
[0017] FIG. 3 shows the turbine blade of FIG. 2, but with a small
piece of the TBC missing and, thus, forming an exposed surface on
the metal substrate on which the high temperature gas can act.
[0018] FIG. 4 shows the turbine blade of FIG. 2, but with a small
hole formed in the metal substrate due to melting of the metal from
exposure to the high temperature gas.
[0019] FIG. 5 shows the turbine blade of FIG. 2, but with a
completed hole formed through the metal substrate and joining the
small cooling air passage, forming an open cooling air passage and
allowing for cooling air to flow through the smaller cooling air
passage and onto the exposed portion of the metal substrate in
order to prevent further melting of the metal substrate.
[0020] FIG. 6 shows a cross-sectional view of the airfoil with a
serpentine internal cooling air passage separated by ribs joining
the pressure side airfoil surface and the suction side airfoil
surface, where the axial direction of the smaller cooling air
passages are aligned with a pulling direction A-E of a multiple
piece mold used to form the airfoil passages.
[0021] FIG. 7 shows an airfoil section with smaller cooling air
passages extending to just under the surface of the airfoil exposed
to high temperature gas, but without the use of a thermal barrier
coating (TBC).
DETAILED DESCRIPTION OF THE INVENTION
[0022] Gas turbine engines include moving blades and stationary
vanes (both considered to be an airfoil) with an internal cooling
passage to direct a cooling fluid (such as air) through the airfoil
for cooling purposes. FIG. 1 shows an airfoil 10 having cooling
passage 20 within the airfoil, and smaller cooling air passages 30
extending from the internal cooling air passage 20 toward an outer
surface 11 of the airfoil 10 on which the high temperature gas is
exposed. These smaller cooling air passages 30 stop short of the
outer surface 11 of the airfoil to a point just under the TBC 40,
thereby forming a closed cooling air passage such that cooling air
from the internal cooling air passage 20 does not flow through the
smaller cooling air passages 30. These cooling air passages 30 can
be referred to as failsafe film cooling holes or breakout passages
or breakout film cooling holes. Also, the failsafe or breakout film
cooling holes can extend just under the surface of the metal
substrate as shown in FIG. 2, or they can extend all the way to the
TBC as in the Barry et al U.S. Pat. No. 6,749,396 B2 described
above.
[0023] The smaller cooling air passages 30 are spaced apart on the
airfoil such that any spallation of the TBC 40 (or, any erosion or
oxidation or other means to remove a portion of the TBC) will
expose one or more of the smaller cooling air passages 30 when the
metal substrate 11 melts away due to exposure to the hot gas
temperature.
[0024] FIG. 2 shows a cross-sectional view of the airfoil with a
TBC 40 on the outer surface 11 of the airfoil 10 on which the high
temperature gas is exposed. The internal cooling air passage 20
forms a closed cooling air passage through the airfoil (although an
open cooling air passage in which cooling air is exhausted into the
working fluid). A plurality of smaller air passages 30 extends from
the internal cooling air passage 20 and toward the outer surface 11
of the airfoil 10 on which the TBC 40 is applied. The smaller
cooling air passages can be used below TBCs or on an airfoil
substrate that does not contain a TBC.
[0025] When a piece of the TBC is removed--for example, such as
spallation, chipping, erosion, and oxidation--the metal substrate
surface 11 of the airfoil 10 exposed to the high temperature gas of
the turbine. Without protection from the TBC, the metal substrate
11 can melt away, damaging the airfoil 10 and therefore reducing
the efficiency of the gas turbine engine. FIG. 3 shows a portion 42
of the TBC removed from the airfoil 10, therefore exposing the
metal substrate 11 to the hot gas.
[0026] With a piece of the TBC missing, the metal substrate is now
exposed to the high temperature gas. The metal substrate begins to
melt away, and eventually will melt a hole 13 to the point in the
airfoil 10 where a small cooling air passage ends as shown in FIG.
4. Thus, the small cooling air passage will open into the melted
hole, forming an open cooling air passage 22 (see FIG. 5). Then,
cooling air from the internal cooling air passage 20 will flow
through the smaller cooling air passage 30 an out of the hole
exposed surface of the airfoil 10, providing cooling air to the
airfoil surface 11 to prevent further damage to the airfoil 10.
[0027] A hollow airfoil like that used in this invention can be
formed by a molding process in which the mold is made up of
multiple pieces joined together at a line 15 forming the midpoint
of the cross-section of the airfoil. FIG. 6 shows an airfoil with a
line 15 along the midpoint of the cross section. The mold pieces
are separated from each other in a respective pulling direction. In
FIG. 6, five pulling directions are show and labeled as A through
E. The pulling direction would be the direction in which the mold
pieces are separated in order to remove the solidified airfoil
after a molding process.
[0028] The smaller cooling passages 30 formed in the airfoil 10 are
directed along an axis parallel to the pulling direction of the
mold piece in which the passage 30 is formed. This makes it easier
to form the smaller cooling air passages 30 in the molding process
and to separate the solidified airfoil from the mold.
[0029] FIG. 7 shows an embodiment of this invention in which no TBC
is used on the airfoil substrate 11. The smaller cooling air
passages 30 are located just below the surface of the airfoil 11.
In this embodiment, the cooling air passing through the internal
cooling air passage 20 cools the airfoil. If this cooling fails on
a portion of the airfoil substrate 11, then a hole will be melted
away such that the smaller cooling air passage 30 is exposed to the
airfoil surface 11 and cooling air flows through the passage 30 to
cool the airfoil and prevent additional melting.
[0030] This invention has been disclosed for use in a blade or vane
of a gas turbine engine. However, areas other than gas turbine
engines can make use of the inventive concept of smaller cooling
air passages located between an internal cooling passage and a
heated surface. When the heated surface is overexposed to the high
temperature, the surface starts to melt away and expose the smaller
cooling air passage to the internal cooling air. Additional melting
away of the substrate is reduced or prevented by the cooling air
flowing through the melted away hole and out through the opening
formed in the melted substrate.
* * * * *