U.S. patent application number 09/818312 was filed with the patent office on 2002-10-03 for turbine blade tip having thermal barrier coating-formed micro cooling channels.
Invention is credited to Darolia, Ramgopal, Lee, Ching-Pang, Schafrik, Robert Edward.
Application Number | 20020141869 09/818312 |
Document ID | / |
Family ID | 25225224 |
Filed Date | 2002-10-03 |
United States Patent
Application |
20020141869 |
Kind Code |
A1 |
Lee, Ching-Pang ; et
al. |
October 3, 2002 |
TURBINE BLADE TIP HAVING THERMAL BARRIER COATING-FORMED MICRO
COOLING CHANNELS
Abstract
The present invention provides for cooling the squealer tip
region of a high pressure turbine blade used in a gas turbine
engine comprising coating the squealer tip with a metallic bond
coat. Micro grooves oriented in the radial direction are fabricated
into the airfoil on the interior surface of the squealer tip above
and substantially perpendicular to the tip cap. A micro groove
oriented in the axial direction is fabricated along the joint
corner between the squealer tip side wall and the tip cap to
connect and act as a plenum with all of the micro grooves oriented
in the radial direction. Tip cap cooling holes are drilled through
the tip cap and connected to the micro groove that ultimately forms
a plenum. TBC ceramic is then deposited on both blade external
surfaces and the tip cavity, forming micro channels from micro
grooves as a result of self shadowing. In this manner, cooling
fluid passes from a cooling fluid source through the tip cap holes
and into the plenum created by the micro channel, subsequently
passing into the micro channels that are oriented in the radial
direction. Cooling fluid is thereby directed through the micro
channels to cool the squealer, exiting in the vicinity of the tip.
Since the TBC is porous, some of the cooling fluid will also flow
through the TBC to provide transpiration cooling. The present
invention further comprises both the cooled blade and squealer tip
region formed by the foregoing methods and the blade and squealer
tip with the micro channels for cooling the squealer tip.
Inventors: |
Lee, Ching-Pang;
(Cincinnati, OH) ; Darolia, Ramgopal; (West
Chester, OH) ; Schafrik, Robert Edward; (Cincinnati,
OH) |
Correspondence
Address: |
GENERAL ELECTRIC COMPANY
ANDREW C HESS
GE AIRCRAFT ENGINES
ONE NEUMANN WAY M/D H17
CINCINNATI
OH
452156301
|
Family ID: |
25225224 |
Appl. No.: |
09/818312 |
Filed: |
March 27, 2001 |
Current U.S.
Class: |
416/97R ;
416/241R |
Current CPC
Class: |
Y02T 50/673 20130101;
F01D 5/187 20130101; F01D 5/20 20130101; F01D 5/288 20130101; F01D
5/183 20130101; Y02T 50/60 20130101; Y02T 50/6765 20180501; Y02T
50/676 20130101; Y10T 29/49318 20150115; Y02T 50/67 20130101 |
Class at
Publication: |
416/97.00R ;
416/241.00R |
International
Class: |
F01D 005/08 |
Claims
What is claimed is:
1. A cooling system for cooling of a squealer tip of a high
pressure turbine blade used in a gas turbine engine comprising: a
superalloy tip cap; a superalloy squealer tip extending outward in
an engine radial direction from the superalloy tip cap into a hot
gas stream of the engine; at least one fluid supply channel having
a first and second end, the first end terminating in a tip cap hole
located on a surface of the tip cap, the second end in fluid
communication with a cooling circuit located within the blade,
wherein the at least one fluid supply channel has a diameter to
permit an effective flow of cooling fluid; a bond coat having a
thickness of about 0.0005" to about 0.005" applied to the tip cap
surface; at least one structured micro channel oriented in a radial
direction formed by shadowing a thermal barrier coating (TBC) onto
a structured micro groove, said micro groove fabricated in a plane
substantially parallel to the plane of the substrate surface in a
generally radial direction on an interior surface of the squealer
tip above and substantially perpendicular to the tip cap in fluid
communication with the at least one fluid supply channel; at least
one structured micro channel oriented in an axial direction formed
by shadowing a TBC onto a structured micro groove fabricated in a
plane substantially parallel to the plane of the substrate surface
at a joint corner between the squealer tip and the tip cap, such
that the structured micro groove oriented in the axial direction
and the resultant axially oriented micro channel is in fluid
communication with the at least one structured micro groove
oriented in the radial direction and resultant axially oriented
micro channel and the tip cap hole; and, a means for providing
cooling fluid to the tip cap hole.
2. The cooling system of claim 1 wherein the at least one of the
structured micro grooves oriented in the radial direction and the
structured micro groove oriented in the axial direction is
fabricated within the bond coat.
3. The cooling system of claim 1 wherein the at least one of the
structured micro grooves oriented in the radial direction and the
structured micro groove oriented in the axial direction is
fabricated into the substrate surface.
4. The cooling system of claim 1 wherein the at least one of the
structured micro grooves oriented in the radial direction and the
structured micro groove oriented in the axial direction is
fabricated within the TBC.
5. The cooling system of claim 1 wherein the at least one of the
structured micro grooves oriented in the radial direction is
fabricated by the process selected from the group consisting of
laser machining, electrochemical machining, electro-discharge
machining and photolithography.
6. The cooling system of claim 1 wherein the TBC is deposited by
shadowing using electron beam physical vapor deposition
(EB-PVD).
7. The cooling system of claim 1 wherein the at least one fluid
supply channel has a diameter of about 0.006" to about 0.020".
8. The cooling system of claim 1 wherein the bond coat has a
thickness of about 0.002".
9. The cooling system of claim 1 wherein the bond coat is an
aluminide selected from the group consisting of NiAl, PtAl and
combinations thereof.
10. The cooling system of claim 1 wherein the bond coat is a
MCrAl(X) where M is an element selected from the group consisting
of Fe, Co and Ni; and X is an element selected from the group
consisting of gamma prime formers, solid solution strengtheners,
grain boundary strengtheners, reactive elements and combinations
thereof.
11. The cooling system of claim 10 wherein X is an element selected
from the group consisting of Zr, Hf, Y and rare earth elements
12. The cooling system of claim 1 wherein the TBC is a porous TBC
and has a thickness of at least about 0.003".
13. The cooling system of claim 1 whereby the cooling fluid is
diffused and flows through the TBC.
14. The cooling system of claim 1 wherein the radial and axial
micro groove size and spacing are about 0.0005" to about
0.010".
15. The cooling system of claim 1 wherein the radial and axial
micro groove size and spacing are about 0.002".
16. The cooling system of claim 1 wherein the cooling fluid is
air.
17. A method for cooling of a squealer tip region of a high
pressure turbine blade used in a gas turbine engine comprising the
steps of: machining at least one fluid supply channel having a
diameter of about 0.006" to about 0.020" in a tip cap of the
turbine blade to allow passage of cooling fluid from a cooling
fluid source within the blade to a surface of the tip cap; applying
a bond coat having a thickness of about 0.0005" to about 0.005" to
the surface of the tip cap and at least one squealer tip wall;
fabricating at least one structured micro groove oriented in the
radial direction in a plane substantially parallel to the plane of
the substrate surface in a generally radial direction on an
interior radially oriented surface of a squealer tip above and
perpendicular to the tip cap in fluid communication with the at
least one fluid supply channel; fabricating at least one structured
micro groove oriented in an axial direction in a plane
substantially parallel to the plane of the substrate surface at a
joint corner between the squealer tip and the tip cap, such that
the structured micro groove oriented in the axial direction is in
fluid communication with the at least one structured micro groove
oriented in the radial direction and the at least one fluid supply
channel; shadowing a TBC to the at least one structured micro
groove oriented in the radial direction and the at least one
structured micro groove oriented in the axial direction to form at
least one radial micro channel and at least one axial micro channel
in fluid communication with each other; and, passing cooling fluid
from the cooling fluid source through the at least one fluid supply
channel into the micro channel oriented in the axial direction, the
axially oriented micro channel supplying cooling fluid to the at
least one micro channel oriented in the radial direction to exit
into the gas stream at the squealer tip.
18. The method of claim 17 wherein the at least one fluid channel
is machined in the tip cap by laser drilling.
19. The method of claim 17 wherein the bond coat is applied a
thickness of about 0.002".
20. The method of claim 17 wherein at least one of the at least one
micro groove oriented in the radial direction and the corresponding
at least one micro channel is fabricated within the bond coat.
21. The method of claim 17 wherein at least one of the at least one
micro groove oriented in the radial direction and the corresponding
at least one micro channel is fabricated in a substrate
surface.
22. The method of claim 17 wherein at least one of the at least one
micro groove oriented in the radial direction and the corresponding
at least one micro channel is fabricated within the TBC.
23. The method of claim 17 wherein the cooling fluid is air.
24. The method of claim 17 wherein the at least one micro groove
oriented in the radial direction and the at least one micro groove
oriented in the axial direction are fabricated by the process
selected from the group consisting of laser machining,
electro-chemical machining, electro-discharge machining and
photolithography.
25. The method of claim 17 wherein the TBC is shadowed using
electron beam physical vapor deposition (EB-PVD).
26. A cooling system for cooling of a squealer tip surface region
of a high pressure turbine blade used in a gas turbine engine
formed by the method of claim 17.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is related to co-pending applications
assigned to the assignee of the present invention which are
identified as Attorney Docket No. 13DV-13803 entitled "Turbine
Airfoil Trailing Edge With Micro Cooling Channels" and 13DV-13804,
entitled "Process for Forming Micro Cooling Channels Inside a
Thermal Barrier Coating System Without Masking Material," and
references co-pending applications assigned to the assignee of the
present invention, which are identified as Attorney Docket No.
13DV-13513 entitled "Directly Cooled Thermal Barrier Coating
System", Attorney Docket No. 13DV-13527 entitled "Multi-layer
Thermal Barrier Coating with Integrated Cooling System," Attorney
Docket No. 13DV-13528 entitled "Integrated Cooling in Thermal
Barrier Coating", and Attorney Docket No. 13DV-13654 entitled
"Cooled Thermal Barrier Coatings On a Turbine Blade Tip," the
contents of which are incorporated herein by reference.
FIELD OF THE INVENTION
[0002] This invention relates generally to gas turbine engines, and
in particular, to a process for cooling a flow path surface region
on a turbine airfoil.
BACKGROUND OF THE INVENTION
[0003] In gas turbine engines, for example, aircraft engines, air
is drawn into the front of the engine, compressed by a
shaft-mounted rotary-type compressor, and mixed with fuel. The
mixture is burned, and the hot exhaust gases are passed through a
turbine mounted on a shaft. The flow of gas turns the turbine,
which turns the shaft and drives the compressor and fan. The hot
exhaust gases flow from the back of the engine, driving it and the
aircraft forward.
[0004] During operation of gas turbine engines, the temperatures of
combustion gases may exceed 3,000.degree. F., considerably higher
than the melting temperatures of the metal parts of the engine aft
of the compressor, which are in contact with these hot gases.
Operation of these engines at gas temperatures that are above the
metal part melting temperatures is a well established art, and
depends in part on supplying a cooling fluid to the outer surfaces
of the metal parts through various methods. Metal parts of these
engines that are particularly subject to high temperatures, and
thus require particular attention with respect to cooling, are, for
example, combustor liners and the metal parts located aft of the
combustor including high pressure turbine airfoils, such as turbine
blades and turbine vanes.
[0005] The hotter the turbine inlet gases, the more efficient is
the operation of the jet engine. There is thus an incentive to
raise the turbine inlet gas temperature. However, the maximum
temperature of the turbine inlet gases is normally limited by the
materials used to fabricate the components downstream of the
combustors such as the vanes and the blades of the turbine. In
current engines, the turbine vanes and blades are made of
nickel-based superalloys, and can operate at temperatures of up to
21000-2200.degree. F. with appropriate well-known cooling
techniques.
[0006] The metal temperatures can be maintained below their melting
levels with current cooling techniques by using a combination of
improved active cooling designs and thermal barrier coatings
(TBCs). For example, with regard to the metal blades and vanes
employed in aircraft engines, some cooling is achieved through
convection by providing passages for flow of cooling air from the
compressor internally within the blades so that heat may be removed
from the metal structure of the blade by the cooling air. Such
blades have intricate serpentine passageways within the structural
metal forming the cooling circuits of the blade.
[0007] Small internal orifices have also been devised to direct
this circulating cooling air directly against certain inner
surfaces of the airfoil to obtain cooling of the inner surface by
impingement of the cooling air against the surface, a process known
as impingement cooling. In addition, an array of small holes
extending from a hollow core through the blade shell can provide
for bleeding cooling air through the blade shell to the outer
surface where a film of such air can protect the blade from direct
contact with the hot gases passing through the engines, a process
known as film cooling.
[0008] In another approach, a TBC is applied to the turbine blade
component, which forms an interface between the metallic component
and the hot gases of combustion. The TBC includes a ceramic coating
that is applied to the external surface of metal parts to impede
the transfer of heat from hot combustion gases to the metal parts,
thus insulating the component from the hot combustion gas. This
permits the combustion gas to be hotter than would otherwise be
possible with the particular material and fabrication process of
the component.
[0009] TBCs include well-known ceramic materials, such as, for
example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do
not adhere well directly to the superalloys used as substrate
materials. Therefore, an additional metallic layer called a bond
coat is placed between the substrate and the TBC. The bond coat may
be made of an overlay alloy, such as a MCrAIX, or other composition
more resistant to environmental damage than the substrate, or
alternatively, the bond coat may be a diffusion nickel aluminide or
platinum aluminide. The surface of the bond coat oxidizes to form a
thin, protective aluminum oxide scale that provides improved
adherence to the ceramic top coatings. The bond coat and overlying
TBC are frequently referred to as a thermal barrier coating
system.
[0010] Multi-layer coatings are well known in the art. For example,
U.S. Pat. No. 5,846,605 to Rickerby et al. is directed to a coating
having a plurality of alternate layers having different structures
that produce a plurality of interfaces. The interfaces provide
paths of increased resistance to heat transfer to reduce thermal
conductivity. Rickerby et al. teaches a traditional bond coat
overlying a metallic substrate bonded to a TBC. The TBC comprises a
plurality of layers, each layer having columnar grains, the
columnar grains in each layer extending substantially perpendicular
to the interface between the bond coating and metallic substrate.
The structure is columnar to ensure that the strain tolerance of
the ceramic TBC is not impaired. The difference in structure of the
layers is the result of variations in the microstructure and/or
density/coarseness of the columnar grains of the ceramic. These
differences assist in providing to resistance to the transfer of
heat across the thermal barrier coating.
[0011] Improved environmental resistance to destructive oxidation
and hot corrosion is desirable. Additionally, the alloying elements
of the bond coat interdiffuse with the substrate alloy at elevated
temperatures of operation, changing the composition of the
protective outer layer. Over time, as the airfoils are refurbished,
walls of the airfoils are consumed, which reduces load carrying
capability and limits blade life. Also, this interdiffusion can
also reduce environmental resistance of the coating, causing loss
of material, as layers of material are lost due to corrosive and
oxidative effects. This interdiffusion and its adverse effects can
be reduced by controlling the temperature of the component in the
region of the bond coat/substrate interface. However, even with the
use of advanced cooling designs and thermal barrier coatings, it is
also desirable to decrease the requirement for cooling; because
reducing the demand for cooling is also well known to improve
overall engine operating efficiency.
[0012] One efficient cooling technique is film cooling. Film
cooling is achieved by passing cooling air through discrete film
cooling holes, typically ranging from 0.015" to about 0.030" in
hole diameters. The film cooling holes are typically drilled with
laser or by electro-discharge machining (EDM) or electro-stem (ES)
machining. Due to mechanical limitations, each film hole has an
angle ranging from 20.degree. to 90.degree. relative to the
external surface. Therefore, each film jet exits from the hole with
a velocity component perpendicular to the surface. But, because of
this vertical velocity component and a complex aerodynamic flow
circulation near the tip of a turbine blade, commonly referred to
as the "squealer tip", each film jet will have a tendency to lift
or blow off from the external surface and mix with the hot exhaust
gases, resulting in poor film cooling effectiveness in the area
surrounding the squealer tip.
[0013] Thus, there is an ongoing need for an improved thermal
barrier coating system surrounding the squealer tip, wherein the
environmental resistance and long-term stability of the thermal
barrier coating system is improved so that higher engine
efficiencies can be achieved. The bond coat temperature limit is
critical to the TBC's life and has an upper limit of about
2100.degree. F. Once the bond coat exceeds this temperature, the
coating system tends to quickly deteriorate, due to high
temperature mechanical deformation and oxidation, as well as from
interdiffusion of elements with the substrate alloy and subsequent
degradation due to loss of its superior environmental resistance.
The coating system can separate from the substrate exposing the
underlying superalloy component to damage from the hot gasses.
[0014] In particular, the squealer tip is the most difficult
location to cool in a turbine blade. The squealer tip is located
away from the convection cooling in the center of the blade, and
the complex aerodynamic flow field near the squealer tip makes film
cooling very inefficient. This inefficient cooling results in tip
deterioration much earlier than desired, and requires tip repairs
after relatively short time in service to recover the tip clearance
for better turbine efficiency.
[0015] As described above, to be effective, a TBC system requires
active cooling on the backside of the location to be cooled. In the
past, a TBC system was not used in the squealer tip region because,
due to physical limitations, no effective backside cooling was
available to make the TBC system effective. During the blade
manufacturing process, to prevent the TBC from adhering to this
area, the squealer tip and tip cap are usually masked during the
TBC coating process. However, recently, in order to reduce
manufacturing costs, the TBC coating application has been extended
to cover the squealer tip region, thus avoiding the time-consuming
masking process. It is therefore advantageous to further improve
squealer tip cooling efficiency by intentionally incorporating an
effective TBC coating system that includes backside cooling to
extend squealer tip life.
[0016] In a co-pending application identified as Attorney Docket
No. 13DV-13654 entitled "Cooled Thermal Barrier Coatings On a
Turbine Blade Tip", channeled apertures were provided to a
substrate in a preselected diameter sufficient to allow passage of
cooling fluid from a cooling fluid source. A sufficiently thick
bond coat was then applied to the substrate such that the bond coat
partially filled the apertures. A porous TBC layer was then applied
to the bond coat to partially occupy the apertures, followed by
application atop the porous TBC layer of an additional dense
ceramic TBC layer, denser than the porous TBC layer overlying the
bond coat . Optionally, a conventional TBC system was applied on
the concave (pressure side) and convex (suction side) of the
airfoil surface.
[0017] In this manner, cooling fluid passes from a cooling fluid
source through a channel aperture substantially adjacent to the
squealer tip into the porous TBC. Because the channel aperture is
at least partially filled with porous TBC material, cooling fluid
flowed through the partially filled passageways. But, due to the
increased resistance to flow, the cooling fluid expanded its flow
path into the adjacent, porous TBC layer, continuing to flow
between the bond coat and the dense coat, thereby directing cooling
fluid to the squealer tip as the fluid exits the blade. However,
because the porosity in the porous TBC layer is randomly
distributed, there is little control of cooling fluid through the
porous TBC and cooling can be irregular. What is needed are further
improved designs that will allow a turbine engine blade squealer
tip to run at higher operating temperatures, thus improving engine
performance without the need for additional cooling air. This can
be achieved by providing a regular cooling pattern through the
porous TBC to the squealer tip. It is also desirable to have a
system that can take advantage of the thermal insulation provided
by TBC. The present invention fulfills this need, and further
provides related advantages.
SUMMARY OF THE INVENTION
[0018] The present invention provides for cooling the squealer tip
region of a high pressure turbine blade used in a gas turbine
engine. The squealer tip is coated with a metallic bond coat and
micro grooves are fabricated in an airfoil oriented in a
substantially radial direction on the interior surface of the
squealer tip above the tip cap. A micro groove oriented in a
substantially axial direction is fabricated along or near the joint
corner between the squealer tip side wall and the tip cap to
provide fluid communication with all of the radial micro grooves.
Tip cap cooling holes are drilled through the tip cap so as to be
in fluid communication with the axial micro groove. TBC ceramic is
then deposited by a shadowing technique on both blade external
surfaces and the tip cavity, forming micro channels from the micro
grooves.
[0019] In this manner, cooling fluid passes from a cooling fluid
source through the tip cap holes and into a plenum created by the
axially-oriented micro channel. The cooling fluid then passes from
the plenum into the radially-oriented micro channels. Cooling fluid
is thereby directed through the micro channels to cool the squealer
tip by exiting to the tip. Since the TBC is porous, some of the
cooling fluid will also provide transpiration cooling as it flows
through the TBC.
[0020] The present invention further comprises both a cooled blade
and a squealer tip region formed by the foregoing methods and
techniques as well as the blade and squealer tip having micro
channels for cooling the squealer tip.
[0021] One advantage of the present invention is that convection
cooling through the micro channels inside or adjacent to a bond
coat layer provides direct and efficient cooling for squealer
tips.
[0022] Another advantage of this invention is that the TBC on the
external surfaces of the squealer tip also becomes a very effective
insulation. This insulating TBC has increased service life due to
the backside cooling provided by the present invention. The
combination of the effective convection cooling inside the micro
channels and effective thermal insulation on the external surfaces
of the squealer tip will lower the temperature of the squealer tip
as compared to conventional and current designs, providing a longer
service life. The beneficial cooling effects of the present
invention create increased adherence of the pressure and suction
side TBC to the airfoil component.
[0023] By removing heat from this region, the integrity of the bond
coat also can be maintained at higher engine firing temperatures by
inhibiting temperature-induced diffusion, resulting in a more
efficient usage of cooling fluid than that of the prior art to
achieve a higher turbine engine efficiency and performance while
improving squealer tip service life.
[0024] Other features and advantages of the present invention will
be apparent from the following more detailed description of the
preferred embodiment, taken in conjunction with the accompanying
figures which illustrate, by way of example, the principles of the
invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0025] FIG. 1 is a representation of a form of current practice
cooling of a turbine blade tip;
[0026] FIG. 2 is a representation of one form of the multi-layer
ceramic coating of the present invention;
[0027] FIG. 3 is a representation of a view at 3-3 of FIG. 2;
[0028] FIG. 4 is a representation of a view at 4-4 of FIG. 2;
and
[0029] FIG. 5 is a representation of a form of the micro grooves
created in the bond coat.
DETAILED DESCRIPTION OF THE INVENTION
[0030] Substrate materials often used in turbine parts or airfoils
for aircraft engines and power generation equipment may include
nickel, cobalt, or iron based superalloys. The alloys may be cast
or wrought superalloys. Examples of such substrates are GTD-111,
GTD-222, Rene' 80, Rene' 41, Rene' 125, Rene' 77, Rene' N4, Rene'
N5, Rene' N6, 4.sup.th generation single crystal superalloy, MX-4,
Hastelloy X, and cobalt-based HS-188. The usage of these materials
at elevated temperatures is limited by their environmental
properties in oxidative and corrosive environments at these
elevated temperatures.
[0031] As shown in FIG. 1, current squealer tip design utilizes
cavity purge holes and pressure side (P/S) film cooling holes that
do not permit cooling fluid to flow into contact with the squealer
tip. In contrast, the present invention provides both convection
cooling and insulation to actively cool the squealer tip.
[0032] Referring now to FIGS. 2 through 5, a bond coat 2 is applied
to the substrate 4 using known techniques to a thickness of from
about 0.0005 inch to about 0.005 inch, preferably about 0.002 inch
in thickness. Bond coat 2 must be of sufficient thickness so that a
structured micro groove 6, 8 (described below) can be formed within
bond coat 2. Bond coat 2 applied to substrate 4 is used for both
increased environmental protection of substrate 4 and to facilitate
adherence of additional layers of ceramic to substrate 4. Bond
coats, such as MCrAl(X)s, are applied to the substrate by, for
example, physical vapor deposition (PVD) processes such as electron
beam evaporation (EB), ion-plasma arc evaporation, sputtering, or
plasma spray processes such as air plasma spray (APS), high
velocity oxy-fuel (HVOF) or low pressure plasma spray (LPPS) and
temperatures can be 1800.degree. F. or higher. PVD processes are
applied in a vacuum and thermal sprays can be controlled so as to
be applied under non-oxidizing conditions. Diffusion aluminide bond
coats may also be formed on substrate in accordance with well-known
techniques.
[0033] For example, bond coat 2 may be a diffusion aluminide, such
as NiAl or PtAl formed by applying aluminum to an appropriate
thickness to a superalloy substrate that may include a thin,
deposited layer of platinum, such as for example by, vapor phase
aluminiding or chemical vapor deposition (CVD). Alternatively, bond
coat 2 may be a MCrAl(X) where M is an element selected from the
group consisting of Fe, Co and Ni and combinations thereof and (X)
is an element selected from the group of gamma prime formers, solid
solution strengtheners, consisting of, for example, Ta, Re and
reactive elements, such as Y, Zr, Hf, Si, and grain boundary
strengtheners consisting of B, C and combinations thereof, in which
deposition typically is by PVD or thermal spray.
[0034] After bond coat 2 has been deposited to a sufficient
thickness, at least one structured radial micro groove 6 is
fabricated within the bond coat 2 substantially parallel to the
bond coat/substrate interface, for example, by laser machining,
electrochemical machining (ECM) electro-discharge machining (EDM)
or photolithography or other process capable of providing the
required sizes and tolerances. The groove size and spacing are
about 0.0005 inch to about 0.010 inch and of such depth that the
groove size does not penetrate the bond coat and expose the
underlying substrate material, preferably about 0.002 inch, with
the radial structured micro groove 6 fabricated in a plane
substantially parallel to the plane of the substrate surface.
[0035] As used herein the term in a generally "radial direction"
refers to a direction extending from the centerline of the engine
outward toward the casing, substantially perpendicular to the flow
of gases through the engine, and "radial surface" or "radial
feature" is a surface or feature that extends parallel to or
substantially parallel to that direction. The term "axial
direction" refers to a direction extending along or substantially
parallel to the centerline of the engine and to the flow of gases
through the engine, and substantially perpendicular to the radial
direction, while the term "axial surface" or "axial feature" is a
surface or feature that extends parallel to or substantially
parallel to that direction.
[0036] Referring to FIGS. 2 and 5, micro grooves 6, extending in a
radial direction are formed on an interior vertical surface 40 of
the squealer tip above the tip cap 11, that is, the surface of the
squealer tip that extends along the airfoil in a radial direction
outward from and substantially perpendicular to tip cap 11 to
connect with that portion of the squealer tip 30 that is
substantially parallel to the surface of the tip cap 11 as shown in
FIG. 2.
[0037] A structured micro groove 8 oriented in an axial direction
is fabricated within the bond coat 2, for example, by laser
machining, ECM, EDM or photolithography. In cross-section,
structured micro groove 8 is identical to micro groove 6, as shown
in FIG. 5. The groove size and spacing are about 0.0005 inch to
about 0.010 inch, preferably about 0.002 inch with the structured
micro groove 8 fabricated into bond coat along a plane
substantially parallel to the plane of the substrate surface, along
the joint corner at the intersection of the radially oriented wall
32 of squealer tip 30 and tip cap 11, such that structured micro
groove 8 oriented in an axial direction forms a plenum or manifold
in fluid communication at one end with structured micro grooves 6
oriented in a radial direction.
[0038] The cross section of the structured micro grooves 6, 8 may
assume any geometric form, for example, a rectangle, a circle, a
triangle or any other shape that will facilitate the flow of
cooling fluid. Structured micro groove 8 oriented in an axial
direction is in fluid communication with a fluid supply source (not
shown) contained within the airfoil component by way of at least
one fluid supply channel machined by known methods through the tip
cap 11. Tip cap hole 10 provides fluid communication micro groove
oriented in an axial direction. These axially oriented micro
grooves form the coolant manifold 17 depicted in FIG. 3, as will be
discussed. Tip cap hole 10 and fluid supply channel are of
sufficient diameter to allow an effective amount of cooling fluid
to pass, preferably about 0.006" to about 0.020" in diameter from
the interior of the airfoil to coolant manifold 17.
[0039] After application of the bond coat to the substrate surface
to a thickness of at least 0.0005 inch, preferably at least about
0.001 inch, and most preferably about 0.002 inch thickness,
structured micro grooves 6, 8 are machined along the radially
oriented wall 32 of squealer tip 30 using known techniques, for
example, by laser machining, electrochemical machining (ECM),
electro-discharge machining (EDM) and photolithographic etching. As
shown in FIG. 4, a TBC 12, preferably a porous TBC, for example,
porous yttria-stabilized zirconia (YSZ), is then applied over bond
coat 2 using well known deposition techniques to achieve a TBC
thickness of at least about 0.003 inch. The porous YSZ structure
can be achieved, for example, by applying the YSZ using a PVD
process such as electron beam PVD or thermal spray processes at
temperatures in the range of 1600.degree.-1800.degree. F., which
are lower than traditional YSZ application temperatures of
1825.degree.-2150.degree. F. Other methods may be utilized
independent of the reduced temperature techniques or in combination
with the reduced temperature techniques to achieve the porous YSZ
structure. Alternatively, the porous TBC can be a ceramic such as a
thin layer of zirconia modified by other refractory oxides such as
oxides formed from Group IV, V and VI elements or oxides modified
by Lanthanide series elements such as La, Nd, Gd, Yb and the
like.
[0040] TBC layer 12 is the subject of a co-pending application
assigned to the assignee of the present invention, which is
identified as Attorney Docket No. 13DV-13528 and titled "Integrated
Cooling in Thermal Barrier Coating". Since the TBC 12 is porous,
some of the cooling fluid will flow through the TBC 12 to provide
transpirational cooling of the TBC layer 12. The porosity of the
TBC layer 12 can be varied as desired using well known methods for
varying deposition densities such as by varying the deposition
temperature.
[0041] As the porous TBC 12 is deposited onto the grooved surfaces,
radial micro channels 14 and axial micro channel 16 are formed
above the grooves 6, 8 due to the shadowing effect of the TBC 12
depositing on top of the peaks 18 (FIG. 5) of the micro grooves 6,
8. The shadowing effect may be best visualized by placing an object
in front of a light source and observing the shadow cast by that
object. Light rays passing around the object would represent TBC
being deposited, while the shadow cast by the object would
represent the void in the deposited TBC. It is this void which
ultimately forms the structured radial micro channel 14 and axial
micro channel 16.
[0042] The dimensions of the radial and axial micro channel 14, 16
in cross section are of sufficient size to allow passage of an
effective amount of cooling fluid, about 0.001 to about 0.010 inch
in diameter, and preferably about 0.002 to 0.004 inch in diameter
when circular, and may be arcuate or may assume other geometric
forms having equivalent dimensions, that is, yielding an equivalent
cross-section size. The cross section of the radial and axial micro
channel 14, 16 may take any preselected form such as, for example,
a parallelogram, rectangle, an oval, a triangle or a circle.
[0043] Cooling fluid (not shown), for example, air, is thereby
routed from an engine cooling fluid supply (not shown) into micro
channel 16 oriented in an axial direction, which serves as a plenum
to supply cooling fluid to at least one micro channel 14 oriented
in a radial direction. After passage through micro channel 14, the
cooling fluid, which is at an elevated temperature, is expelled,
typically into the gas stream at the squealer tip axial surface,
although, as mentioned above, there is some transpiration through
the TBC. In this manner, the bond coat 2 is kept at a reduced
temperature through active convection cooling throughout the entire
squealer tip.
[0044] In another embodiment of the present invention, the
structured micro grooves described above first are partially
machined, for example, by a laser, by an ECM technique, cast during
manufacture, or etched, into the surface of the turbine blade tip
substrate. A bond coat is then applied to the substrate so as to
coat, but not fill, the structured micro grooves. A porous TBC
layer is deposited using a shadowing technique, as described above,
to form the radial and axial micro channels at the interface
between the bond coat and the substrate. At least one micro groove
is in communication with the cooling fluid supply to provide active
cooling of the squealer tip in the manner as described above.
[0045] In still another embodiment of the present invention, a bond
coat and first layer of TBC is applied to the relatively smooth
substrate comprising at least the vertical walls of the squealer
tip adjacent the tip cap and the tip cap using known techniques.
The bond coat may also be applied to the entire squealer tip if no
masking operations are to be performed on the squealer tip. The
effect of applying bond coat to this area, and any subsequent TBC,
is that the applied material will be worn off by the engine during
its initial cycles, as the airfoils "rub in." If necessary, at
least one hole is machined for example, using a laser, through the
bond coat and first layer of TBC to communicate with a cooling
fluid supply, to allow for flow of cooling fluid as previously set
forth. Structured radial and axial micro grooves are machined in
the first TBC layer such as by laser machining. A second TBC layer
is then applied using the shadowing technique as described above.
In this manner, the location of the structured micro groove, hence,
the micro channel, may be placed at any preselected position within
the TBC layer between the bond coat and the hot gaseous
atmosphere.
[0046] Although the present invention has been described in
connection with specific examples and embodiments, those skilled in
the art will recognize that the present invention is capable of
other variations and modifications within its scope. These examples
and embodiments are intended as typical of, rather than in any way
limiting on, the scope of the present invention as presented in the
appended claims.
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