U.S. patent application number 10/954567 was filed with the patent office on 2006-04-06 for gas turbine airfoil film cooling hole.
This patent application is currently assigned to ALSTOM Technology Ltd.. Invention is credited to George Liang.
Application Number | 20060073015 10/954567 |
Document ID | / |
Family ID | 36125729 |
Filed Date | 2006-04-06 |
United States Patent
Application |
20060073015 |
Kind Code |
A1 |
Liang; George |
April 6, 2006 |
Gas turbine airfoil film cooling hole
Abstract
A gas turbine airfoil (1) with a pressure sidewall and a suction
sidewall (6) comprises several internal cooling passages, through
which cooling air flows, and several film cooling holes (7) that
extend from the internal cooling passages to the outer surface (11)
of the suction sidewall (6). According to the invention the film
cooling holes (7) are oriented in the radial outward direction as
well as toward the trailing edge of the airfoil (1). They each
comprise a diffused section with sidewalls that are diffused in the
radial outward direction, in the radial inward direction as well as
toward downstream direction. Cooling air flowing through the
diffused film cooling hole (7) onto the outer surface (6) of the
airfoil (1) forms an air curtain preventing plugging of the film
cooling hole by particles entrained in the hot gas flow.
Inventors: |
Liang; George; (Palm City,
FL) |
Correspondence
Address: |
BUCHANAN INGERSOLL PC;(INCLUDING BURNS, DOANE, SWECKER & MATHIS)
POST OFFICE BOX 1404
ALEXANDRIA
VA
22313-1404
US
|
Assignee: |
ALSTOM Technology Ltd.
Baden
CH
|
Family ID: |
36125729 |
Appl. No.: |
10/954567 |
Filed: |
October 1, 2004 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F05D 2250/324 20130101;
F05D 2250/314 20130101; F01D 5/186 20130101 |
Class at
Publication: |
416/097.00R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. Gas turbine airfoil with a pressure sidewall and a suction
sidewall extending from a root to a tip and from a leading edge to
a trailing edge of the airfoil comprises several internal cooling
passages, through which cooling air flows, and several film cooling
holes that extend from the internal cooling passages to the outer
surface of the suction sidewall wherein the film cooling holes each
comprise a metering section of cylindrical shape and a diffused
section with sidewalls that are diffused in the radial direction
and in the streamline direction both with respect to the
longitudinal axis of the film cooling hole.
2. Gas turbine airfoil according to claim 1 wherein the sidewalls
of the film cooling holes are diffused in the radial outward
direction toward the tip of the airfoil and in the radial inward
direction toward the root of the airfoil.
3. Gas turbine airfoil according to claim 2 wherein the sidewall
closest to the tip of the airfoil is diffused with respect to the
longitudinal axis of the film cooling hole in the radially outward
direction by an angle that is in the range of 3 to 7.degree.,
preferably about 5.degree., and that the sidewall closest to the
root of the airfoil is diffused with respect to the longitudinal
axis of the film cooling hole in the radially inward direction by
an angle that is in the range of 7 to 12.degree., preferably about
10.degree..
4. Gas turbine airfoil according to claim 3 wherein the axis of the
film cooling hole is oriented at an angle with respect to the
streamwise direction that is in the range of 45 to 55.degree.,
preferably about 50.degree., where the streamwise direction is
perpendicular to the radial direction and perpendicularly away from
the outer surface of the suction side of the airfoil
5. Gas turbine airfoil according to claim 4 wherein the
longitudinal axis of the film cooling hole is oriented at an angle
with respect to the downstream direction and toward the trailing
edge of the airfoil, where the angle .epsilon. is in the range of
35 to 45.degree. and preferably about 40.degree..
6. Gas turbine airfoil according to claim 5 wherein the sidewall of
the film cooling hole that is closer to the trailing edge of the
airfoil is diffused at an angle with respect to the longitudinal
axis of the film cooling hole that is in the range of 7 to
12.degree. and preferably about 10.degree..
7. Gas turbine airfoil according to claim 1 wherein the film
cooling holes with diffused sidewalls are positioned in one or more
radially extending rows on the suction side of the airfoil.
8. Gas turbine airfoil according to claim 7 wherein some rows of
film cooling holes are placed in the strike-zone of the airfoil
where heavy particles primarily strike the suction side of the
airfoil.
9. Gas turbine airfoil according to claim 8 wherein the film
coverage for the rows of film cooling holes in strike-zone is at
least 75%.
Description
FIELD OF INVENTION
[0001] This invention pertains to gas turbine airfoils and in
particular to a cooling construction with film cooling holes and
the prevention of contamination and plugging of the film cooling
holes.
BACKGROUND ART
[0002] The airfoils of gas turbines, turbine rotor blades and
stator vanes, require extensive cooling in order to keep the metal
temperature below a certain allowable level and prevent damage due
to overheating from the hot gas flow. Typically, such airfoils are
designed with hollow spaces and a plurality of passages and
cavities within the airfoil for cooling fluid to flow through. The
cooling fluid is typically air bled from the compressor having a
higher pressure and lower temperature compared to the gas traveling
through the turbine. The higher pressure forces the air through the
cavities and passages as it transports the heat away from the
airfoil walls. The cooling construction further comprises film
cooling holes leading from the hollow spaces within the airfoil to
the external surfaces of the leading and trailing edge as well as
to the suction and pressure sidewalls. The cooling fluid flows
through the film cooling holes to the airfoil outer surface and
flows along the outer surfaces forming a film of cooling air of a
given penetration depth.
[0003] One problem encountered in the design of the gas turbine
airfoils is caused by particles entrained in the hot gas flow.
During engine operation small particles remain entrained in the gas
flow while larger and heavier particles impinge on the airfoil
surface and can cause local impact damage. This problem is
encountered in particular on the suction side of the airfoil and
downstream from the leading edge. The particles can furthermore
plug the film cooling holes especially on the suction side. When
the particles strike the surface of the airfoil they solidify on
the cooled airfoil wall and accumulate between the holes and
eventually plug up the holes.
[0004] The particles hence not only cause impact damage but can
also prevent the air from reaching the airfoil outer surface and
cause secondary damage due to loss of film cooling and resultant
overheating of the airfoil. Plugging of the film cooling holes is
particularly likely if the film cooling holes are oriented in the
axial direction as the particles travel in the axial as well as
radial direction due to centrifugal forces.
[0005] U.S. Pat. No. 5,688,104 discloses an airfoil with a cooling
construction comprising internal cavities for the cooling air to
flow and film cooling holes with a metering section and a diffusing
section. FIG. 4 of the patent disclosure shows an axial diffusion
hole with a diffusing angle with respect to the hole axis and in
the plane of the shown cross-section that is approximately parallel
to the root of the airfoil.
SUMMARY OF INVENTION
[0006] It is an object of the invention to provide a cooling
construction for the suction side of a gas turbine airfoil that
prevents contamination and plugging of film cooling holes by hot
particles entrained in the gas flow.
[0007] According to the invention a gas turbine airfoil with a
pressure sidewall and a suction sidewall extending from a root to a
tip and from a leading edge to trailing edge comprises several
internal cooling passages within through which cooling air can flow
and cool the airfoil. Several film cooling holes lead from the
internal cooling passages to the outer surfaces on the suction
sidewall providing cooling air to flow onto the airfoil outer
surface. The film cooling holes on the suction side of the airfoil
each comprise a metering section of cylindrical shape and a
diffused section. The sidewalls of the diffused section are
diffused with respect to the film cooling hole axis in the radial
direction as well as in the downstream direction toward the
trailing edge of the airfoil.
[0008] The film cooling hole according to the invention causes the
cooling air flow exiting from the film cooling holes onto the
surface of the suction sidewall to diffuse in the radial as well as
in the downstream direction. The downstream direction is defined
here as the direction along the tangent of the airfoil at the point
of the exit port of the film cooling hole and in the plane
perpendicular to the radial direction. The diffusion in the radial
as well as in the downstream direction brings about an improved
film coverage, which seals the airfoil surface from hot particles
in the manner of an air curtain.
[0009] Particles in the gas stream that approach a film cooling
hole will flow around or over this air curtain. Thus the film
cooling hole is protected from particles accumulating in its
vicinity and contamination and plugging of the film cooling hole by
such particles is prevented.
[0010] In a particular embodiment of the invention the sidewalls of
the film cooling holes are diffused in the radially outward
direction toward the tip of the airfoil as well as in the radially
inward direction toward the root of the airfoil. This further
improves film coverage between neighboring film cooling hole exit
ports in the radial direction.
[0011] In a further particular embodiment of the invention the
diffusion angle of the sidewall with respect to the film cooling
hole axis and in the radially outward direction toward the tip of
the airfoil is in the range of 3 to 70, preferably about 5.degree..
This sidewall is diffused with respect to the film cooling hole
axis and in the radially inward direction toward the root of the
airfoil by an angle in the range of 7 to 12.degree., preferably
about 10.degree..
[0012] In a preferred embodiment of the invention the axis of the
film cooling hole is at an angle with respect to the streamwise
direction that is in the range of 45 to 55.degree. and preferably
about 50.degree.. The streamwise direction is defined as the
direction perpendicularly away from the outer surface of the
suction side, perpendicular to the radial direction.
[0013] The film cooling axis is furthermore oriented at an angle
with respect to the downstream direction that is in the range of 35
to 45.degree. and preferably about 40.degree.0. The downstream
direction is here defined as the direction along the tangent to the
suction side at the point of the exit port of the film cooling hole
and pointing away from the leading edge and toward the trailing
edge of the airfoil.
[0014] In a further particular embodiment of the invention the
sidewall of the film cooling hole is diffused with respect to the
film cooling axis and in the downstream direction toward the
trailing edge of the airfoil by an angle that is in the range of 7
to 12.degree., preferably about 10.degree..
[0015] In a further embodiment of the invention the film cooling
holes are arranged in rows extending radially from the root to the
tip of the airfoil. Some of the rows are arranged in the impact
zone or so-called strike-zone of the airfoil, which can be
approximated by knowing the wheel speed and hot gas axial
velocity.
[0016] In a further preferred embodiment of the invention the film
coverage in the rows of film cooling holes in the strike zone is at
least 75%. That is, the ratio of the film hole break-out length to
the filk hole spacing is at least 0.75.
BRIEF DESCRIPTION OF THE FIGURES
[0017] FIG. 1 shows view of a gas turbine airfoil and its suction
side with several rows of film cooling holes and an enlarged view
of the film cooling holes on the suction side of the airfoil,
[0018] FIG. 1a shows an enlarged view of the exit ports of the film
cooling holes on the suction side of the airfoil,
[0019] FIG. 2 shows a cross-section of the airfoil along line II-II
and the film cooling holes with their diffused sidewalls in the
radial direction,
[0020] FIG. 3 a cross-section of the airfoil along line III-III and
an film cooling hole with the diffused sidewall in the streamline
direction,
[0021] FIG. 3a a detailed view of the film cooling hole in the
cross-section along line III-III,
[0022] FIG. 4 shows a perspective of an individual film cooling
hole with diffused sidewalls in the radial and streamline
directions.
DETAILED DESCRIPTION OF THE INVENTION
[0023] FIG. 1 shows a typical airfoil 1 of a gas turbine extending
from a root 2 to a tip 3 and frown a leading edge 4 to a trailing
edge 5. The figure shows the airfoil from its suction side 6.
Several film cooling holes 7 are arranged In radially extending
rows 8a, 8b, and 8c on the suction side 6. The film cooling holes 7
are realized with diffused sidewalls resulting in an irregularly
shaped exit port 9 as shown in an enlarged view in FIG. 1a.
[0024] The film cooling holes extend from internal cooling passages
within the airfoil 1 through the suction sidewall 6 to its outer
surface 11. They provide convective cooling of the sidewall from
within as well as film cooling of the sidewall outer surface
11.
[0025] In one embodiment of the invention the film cooling holes
with diffused sidewalls are arranged in one or more radially
extending rows on the suction side of the airfoil.
[0026] The number of rows of film cooling holes and their position
is determined based on the metal temperature required for the
particular airfoil. Some of the rows of film cooling are also
placed in the so-called strike zone of the suction side of the
airfoil where the particles and in particular the larger particles
of approximately 40 mils (approximately 1 mm) diameter, strike the
airfoil's suction side. The smaller particles, with a size of 2
mils (approximately 0.05 mm) for example, tend to travel along with
the hot gas flow. The larger particles however, impact the airfoil
suction side surface downstream of the leading edge and in
particular in the radially outer half of the airfoil due to
centrifugal force (beyond the 50% span).
[0027] The arrangement of the film cooling holes in rows brings
about an improved film coverage of this strike zone and prevents
heavy particles from damaging the airfoil. The particular design of
the film cooling hole with its orientation of the film cooling hole
axis and the diffusion of the hole sidewalls result in an air
curtain that prevents the plugging of the holes just in this
zone.
[0028] The rows 8a and 8b are positioned in the so-called
strike-zone of the airfoil where the larger and heavier particles
entrained in the hot gas flow typically strike the airfoil. Row 8c
is placed far from this strike-zone near the trailing edge of the
airfoil.
[0029] In the rows 8a and 8b the ratio of the film hole break-out
length to the film hole spacing is preferably at least 0.75.
[0030] FIG. 2 shows a cross-section of the suction sidewall 6 of
the airfoil with three of the several film cooling holes 7. They
extend from an internal cooling space 10 through the sidewall 6 to
the outer surface 11 of the sidewall. The film cooling holes 7 each
comprise a first metering section 12 with cooling hole sidewalls 13
that run parallel to the film cooling hole axis 14. They further
comprise a second, diffused section 15 designed according to the
invention to create an air curtain on the surface 11 suction
sidewall and to prevent plugging of the film cooling hole by hot
particles. The diffused section 15 has a sidewall 16a that is
diffused with respect to the film cooling hole longitudinal axis 14
in the outward radial direction A toward the tip of the airfoil by
an angle .alpha.. This angle .alpha. is in the range of 3 to
7.degree.. A sidewall 16b of the film cooling hole diffused section
15 that is closer to the root of the airfoil is oriented at angle
.beta. toward the root of the airfoil in the radially inward
direction B with respect to the film cooling hole axis 13. This
angle .beta. is in the range of 7 to 12.degree.. Furthermore, the
axis 14 of the film cooling hole is oriented at an angle .gamma.
with respect to the direction C, which is the steamline direction.
The angle .gamma. is in the range of 45 to 55.degree.. Since a
heavy particle is generally observed to occur in the outboard 50%
airfoil span, the particle trajectory angles can be estimated to be
in the range of 45 to 55.degree. with respect to the direction C.
The preferable angle for .alpha. is 5.degree. outward from the film
hole axis 14, the preferable angle for .beta. is 10.degree. inward
from the film hole axis 14, and the preferable angle for .gamma. is
50.degree.. With film cooling holes of such dimensions an air
curtain is formed in the range of 40 to 55.degree. with respect to
the direction C, which is equal or greater than the hot particle
trajectory angles.
[0031] FIGS. 3 and 3a show a further cross-section of the gas
turbine airfoil 1 perpendicular to the cross-section of FIG. 2 and
illustrates the diffusion of a sidewall 16c of the film cooling
hole 7 on the suction side 6 in the plane shown. Again the
longitudinal axis 14 of the film cooling hole is shown. The
sidewall 16c closer to the trailing edge 5 of the airfoil 1 is
diffused with respect to the film cooling hole axis 14 at an angle
.delta. that is in the range of 7 to 12.degree. and preferably
about 10.degree.. The axis 14 of the film cooling hole is oriented
at an angle .epsilon. with respect to the streamline direction D
and toward the trailing edge 5. This angle .epsilon. is in the
range of 35 to 45.degree., preferably 40.degree.. The streamline
direction D follows the tangent to the airfoil 1 at the point of
the exit port of the film cooling hole and in the plane of FIG.
3.
[0032] FIG. 4 shows for a better understanding of the
"multiple-diffusion" film cooling hole, a perspective view of the
hole. It shows the straight-walled metering section and the
diffused section with the exit port 9. The diffused sidewalls 16a,
16b are shown with the diffusion angles .alpha., .beta., with
respect to the film cooling hole axis 14.
TERMS USED IN THE FIGURES
[0033] 1 gas turbine airfoil [0034] 2 root of the airfoil [0035] 3
tip of the airfoil [0036] 4 leading edge [0037] 5 trailing edge
[0038] 6 suction side [0039] 7 film cooling hole [0040] 8a,8b,8c
rows of film cooling holes [0041] 9 exit port of film cooling hole
[0042] 10 internal cooling space [0043] 11 outer surface of suction
side [0044] 12 metering section [0045] 13 film cooling hole
sidewall in metering section [0046] 14 film cooling hole
longitudinal axis [0047] 15 diffused section [0048] 16a sidewall
diffused in outward radial direction [0049] 16b sidewall diffused
in inward radial direction [0050] 16c sidewall diffused in
streamline direction [0051] .alpha. angle of diffusion in outward
radial direction [0052] .beta. angle of diffusion in inward radial
direction [0053] .gamma. angle of orientation of longitudinal axis
[0054] .delta. angle of diffusion in streamline direction [0055]
.epsilon. angle of orientation of longitudinal axis in streamline
direction [0056] A radial outward direction [0057] B radial inward
direction [0058] C streamline direction [0059] D downstream
direction
* * * * *