U.S. patent application number 13/075234 was filed with the patent office on 2011-09-29 for blade for a gas turbine.
This patent application is currently assigned to ALSTOM TECHNOLOGY LTD. Invention is credited to Erich Kreiselmaier, Jose McFeat Anguisola, Christoph Nagler, Sergei Riazantsev.
Application Number | 20110236223 13/075234 |
Document ID | / |
Family ID | 40130846 |
Filed Date | 2011-09-29 |
United States Patent
Application |
20110236223 |
Kind Code |
A1 |
Nagler; Christoph ; et
al. |
September 29, 2011 |
BLADE FOR A GAS TURBINE
Abstract
A blade for a gas turbine includes an airfoil extending in a
longitudinal direction and extending transversely to the
longitudinal direction between a leading edge and a trailing edge.
The airfoil has a pressure side, a suction side, and a slot-like
cooling medium outlet extending along the trailing edge. The
cooling medium outlet is configured to discharge cooling medium
supplied from an inner space of the blade. A platform extends
transversely to the longitudinal direction. An end of the airfoil
merges into an underside of the platform and has a transition from
the airfoil to the platform at the trailing edge of the airfoil
that increases in thickness in a direction toward the underside of
the platform. The cooling medium outlet extends into the platform
so as to reduce an operating temperature in a region of the
transition from the blade airfoil to the platform.
Inventors: |
Nagler; Christoph; (Zurich,
CH) ; Kreiselmaier; Erich; (Stetten, CH) ;
McFeat Anguisola; Jose; (Lauchringen, DE) ;
Riazantsev; Sergei; (Nussbaumen, CH) |
Assignee: |
ALSTOM TECHNOLOGY LTD
Baden
CH
|
Family ID: |
40130846 |
Appl. No.: |
13/075234 |
Filed: |
March 30, 2011 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
PCT/EP2009/062213 |
Sep 21, 2009 |
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13075234 |
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Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F01D 5/187 20130101;
F05D 2240/80 20130101; Y02T 50/676 20130101; F05D 2250/24 20130101;
Y02T 50/60 20130101 |
Class at
Publication: |
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Foreign Application Data
Date |
Code |
Application Number |
Sep 30, 2008 |
CH |
01548/08 |
Claims
1. A blade for a gas turbine, the blade comprising: an airfoil
extending in a longitudinal direction and extending transversely to
the longitudinal direction between a leading edge and a trailing
edge, the airfoil having a pressure side, a suction side, and a
slot-like cooling medium outlet extending along the trailing edge,
the outlet configured to discharge cooling medium supplied from an
inner space of the blade; a platform extending transversely to the
longitudinal direction, a first end of the airfoil merging into an
underside of the platform and having a transition from the airfoil
to the platform at the trailing edge of the airfoil that increases
in thickness in a direction toward the underside of the platform,
and wherein the cooling medium outlet extends into the platform so
as to reduce an operating temperature in a region of the transition
from the blade airfoil to the platform.
2. The blade as recited in claim 1, wherein the transition from the
blade airfoil to the platform has a thickness profile having a
substantially exponential shape corresponding to at least one of an
inverted rampant pyramid, an inverted truncated pyramid, and an
inverted virtual pyramid.
3. The blade as recited in claim 1, wherein the transition from the
blade airfoil to the platform has an approximately elliptical
profile.
4. The blade as recited in claim 1, wherein the transition from the
blade airfoil to the platform extends to an edge of the
platform.
5. The blade as recited in claim 2, wherein the transition from the
blade airfoil to the platform extends to an edge of the
platform.
6. The blade as recited in claim 3, wherein the transition from the
blade airfoil to the platform extends to an edge of the
platform.
7. The blade as recited in claim 1, wherein the cooling medium
outlet is formed between a pressure-side wall of the blade airfoil
and a suction-side wall of the blade airfoil, and wherein the
transition from the blade airfoil to the platform along the
pressure-side wall has a curvilinear edge profile such that a wall
thickness of the pressure-side wall in a region of the transition
is approximately equal to a wall thickness in a remaining region of
the blade airfoil.
8. The blade as recited in claim 7, wherein the transition from the
blade airfoil to the platform extends to an edge of the platform.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
[0001] This application claims priority to International Patent
Application No. PCT/EP2009/062213, filed Sep. 21, 2009, which
claims priority from Swiss Patent Application No. 01548/08, filed
Sep. 30, 2008, each of which are incorporated by reference herein
in their entirety. The International Application was published as
WO2010/037659 on Apr. 8, 2010.
FIELD
[0002] The present invention relates to the field of gas turbines,
and particularly relates to a blade for a gas turbine.
BACKGROUND
[0003] The requirements for increasing the efficiency of gas
turbines leads to the thickness at the trailing edges of the blade
airfoils of the blades which are fitted in the gas turbines having
to be continuously reduced. This results in a geometry of the blade
as is exemplarily shown in cross section in FIG. 1. The blade 10 of
FIG. 1 extends in the manner of an airfoil profile transversely to
its longitudinal direction between a rounded leading edge 15 and a
comparatively sharply tapering trailing edge 16. The blade 10 has a
(concave) pressure side 13 and a (convex) suction side 14 with
corresponding walls 13' and 14'. A gaseous cooling medium or
coolant is fed in the hollow inner space 17 and discharged into the
environment inter alia through a cooling medium outlet which is
formed at the trailing edge 16. A particularly sharply tapering,
slender trailing edge 16 in this case is achieved by the cooling
medium outlet 18 being arranged entirely on the pressure side 13 of
the blade 10, and by the two walls 13' and 14' being constructed
especially thin in the region of the trailing edge 16.
[0004] If, as is shown in the perspective view of FIG. 2, the blade
airfoil 11 at the end of its extent merges in the longitudinal
direction into a platform 12 which lies transversely to the
longitudinal direction and is delimited by this platform 12, the
transition of the blade airfoil 11 to this platform 12 in the
region of the trailing edge 16 represents a typical factor which
limits the service life of a gas turbine component because it is
exposed to superposition of high thermal stress which is brought
about by the thermo-mechanical mismatch between platform 12 and
blade airfoil 11, and to mechanical stress peaks which are brought
about by loading of the blades as a result of the gas flow.
Reducing the thickness of the trailing edge 16 causes an increase
of the stress in this critical region so that when designing the
blade measures have to be considered in order to achieve and to
ensure a sufficiently long service life.
SUMMARY
[0005] An aspect of the present invention is to further develop a
blade for a gas turbine so that despite a low thickness at the
trailing edge of the blade airfoil a satisfactory service life is
achieved.
[0006] In an embodiment, the present invention provides a blade for
a gas turbine including an airfoil extending in a longitudinal
direction and extending transversely to the longitudinal direction
between a leading edge and a trailing edge. The airfoil has a
pressure side, a suction side, and a slot-like cooling medium
outlet extending along the trailing edge. The cooling medium outlet
is configured to discharge cooling medium supplied from an inner
space of the blade. A platform extends transversely to the
longitudinal direction. An end of the airfoil merges into an
underside of the platform and has a transition from the airfoil to
the platform at the trailing edge of the airfoil that increases in
thickness in a direction toward the underside of the platform. The
cooling medium outlet extends into the platform so as to reduce an
operating temperature in a region of the transition from the blade
airfoil to the platform.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] Exemplary embodiments of the present invention are described
in more detail below, with reference to the drawings. Some
non-essential elements of the invention have been omitted. Like
elements are provided with the same designations in the different
figures. In the drawings:
[0008] FIG. 1 shows a simplified cross section through an gas
turbine blade with a narrow trailing edge and a cooling medium
outlet;
[0009] FIG. 2 shows a sharp transition between blade airfoil and
platform in a blade such as that shown in FIG. 1; and
[0010] FIG. 3 shows a low-stress transition between blade airfoil
and platform according to an exemplary embodiment of the
invention.
DETAILED DESCRIPTION
[0011] In an embodiment, the present invention provides a
transition from the blade airfoil to the platform at the trailing
edge that has a transition thickness profile in which the thickness
increases above average the closer it gets to the underside of the
platform, and that the cooling medium outlet is extended right into
the platform for reducing the temperature in the region of the
transition from the blade airfoil to the platform. As a result of
increasing the thickness of the trailing edge towards the platform
the mechanical stress in the transition region is reliably reduced.
Extending the cooling medium outlet right into the platform leads
to improved cooling there so that thermally induced stresses are
also significantly reduced.
[0012] In one embodiment, the transition thickness profile has an
essentially exponential shape which resembles an inverted, very
slender, truncated pyramid or an inverted virtual pyramid. As a
result of this an especially "smooth" transition between trailing
edge and platform is achieved.
[0013] In another embodiment, the transition from the blade airfoil
to the platform has an approximately elliptical transition border
profile which also reduces stresses.
[0014] Furthermore, it is advantageous if according to another
development the trailing edge at the transition from the blade
airfoil to the platform is extended up to the edge of the
platform.
[0015] In an other embodiment, the cooling medium outlet is formed
between a pressure-side wall of the blade airfoil and a
suction-side wall of the blade airfoil, and in that the
pressure-side wall has a curvilinear transition edge profile in the
transition from the blade airfoil to the platform in such a way
that the wall thickness of the pressure-side wall in the region of
the transition from the blade airfoil to the platform is
approximately equal to the wall thickness in the remaining region
of the blade airfoil.
[0016] In FIG. 3, a blade 20 for a gas turbine with a low-stress
transition between blade airfoil 11 and platform 12 according to an
exemplary embodiment of the invention is reproduced. The blade 20
of the exemplary embodiment comprises a blade airfoil 11 which
extends in a longitudinal direction and in the manner of a wing
extends transversely to the longitudinal direction between a
leading edge 15 and a trailing edge 16, and has a pressure side 13
and a suction side 14. At the upper (or lower) end the blade
airfoil 11 merges into a platform 12 which lies transversely to the
longitudinal direction and projects laterally across the blade
cross section. A slot-like cooling medium outlet 18 which extends
along the trailing edge 16 is provided at the trailing edge 16 of
the blade airfoil 11, through which a cooling medium, for example
cooling air, which is fed via the (hollow) inner space 17 of the
blade 20, is discharged. The trailing edge 16 with its thin walls
13' and 14' is very narrow in construction. In order to reduce the
thermal stresses at the transition between the narrow trailing edge
16 and the solid platform 12, according to embodiments of the
invention the transition has a transition thickness profile 21 in
which the thickness D increases above average the closer it gets to
the underside 12' of the platform 12. At the same time, the cooling
medium outlet 18 is extended (extension 19) right into the platform
12 for reducing the local temperature in the region of the
transition from the blade airfoil 11 to the platform 12.
[0017] The transition thickness profile 21 has an essentially
exponential shape and as a result resembles an inverted rampant
pyramid. It is especially favorable in this case with regard to the
stress distribution if the transition from the blade airfoil 11 to
the platform 12 has an approximately elliptical transition border
profile 22. While in the case of the conventional blade according
to FIG. 2 the trailing edge 16 of the blade airfoil 11 terminates
inside the platform 12 and does not extend as far as the boundary
of the platform 12, in the case of the exemplary embodiment of FIG.
3 the trailing edge 16 at the transition from the blade airfoil 11
to the platform 12 is extended up to the edge 12'' of the platform
12.
[0018] As is to be seen in FIG. 3, the cooling medium outlet 18 is
delimited by the pressure-side wall 13' and the suction-side wall
14' of the blade airfoil 11. The pressure-side wall 13' in this
case has a curvilinear transition edge profile 23 in the transition
from the blade airfoil 11 to the platform 12 in such a way that the
wall thickness of the pressure-side wall 13' in the region of the
transition from the blade airfoil 11 to the platform 12 is
approximately equal to the wall thickness in the remaining region
of the blade airfoil 11.
[0019] In all, an appreciable improvement of the service life at
the transition between the blade-airfoil trailing edge and the
platform of a gas turbine blade is achieved by the invention as a
result of the following measures: [0020] (1) Extending the cooling
medium outlet (cooling slot) into the platform in order to reduce
the metal temperature in the critical region by feeding cooling
medium, wherein convective cooling of the walls on both sides takes
place. [0021] (2) Relocating the blade-airfoil trailing edge to the
boundary of the platform in order to lower stress and to make the
design of the blade independent of deviations in the radial
position of the cast core. [0022] (3) Introducing a transition
thickness profile of a rampant pyramid type by increasing the
height and introducing a special elliptical contour of the fillet
at the transition between blade airfoil and platform in the region
of the trailing edge. [0023] (4) Introducing a specially
curvilinear transition edge profile towards the pressure side of
the blade airfoil on the fillet at the transition between blade
airfoil and platform in the region of the trailing edge in order to
achieve a wall thickness in the transition region which corresponds
to the wall thickness of the blade airfoil, as a result of which
stress and metal temperature is reduced and service life at the
transition is increased.
List of Designations
[0024] 10, 20 Blade (gas turbine)
[0025] 11 Blade airfoil
[0026] 12 Platform [0027] 12' Underside (platform)
[0028] 12'' Edge (platform)
[0029] 13 Pressure side
[0030] 13' Wall (pressure side)
[0031] 14 Suction side
[0032] 14' Wall (suction side)
[0033] 15 Leading edge
[0034] 16 Trailing edge
[0035] 17 Inner space
[0036] 18 Cooling medium outlet (slot-like)
[0037] 19 Extension (cooling medium outlet)
[0038] 21 Transition thickness profile
[0039] 22 Transition border profile
[0040] 23 Transition edge profile
[0041] D Thickness (transition thickness profile)
* * * * *