U.S. patent application number 12/892555 was filed with the patent office on 2011-02-17 for blade for a rotating thermal machine.
This patent application is currently assigned to ALSTOM Technology Ltd. Invention is credited to Willy Heinz HOFMANN, Michael HUBER.
Application Number | 20110038733 12/892555 |
Document ID | / |
Family ID | 39870460 |
Filed Date | 2011-02-17 |
United States Patent
Application |
20110038733 |
Kind Code |
A1 |
HOFMANN; Willy Heinz ; et
al. |
February 17, 2011 |
BLADE FOR A ROTATING THERMAL MACHINE
Abstract
A blade for a rotating thermal machine, such as a stator blade
for a low-pressure turbine of a gas turbine with sequential
combustion, includes a blade airfoil which extends essentially in
the radial direction, is exposed to circumflow by an operating
medium in a flow direction, and is delimited by a leading edge and
a trailing edge in the flow direction. The concurring fluidic and
constructional demands on the blade with sharp inclination of the
flow are met by the blade airfoil being structured in such a way
that an angle which the flow lines form with the trailing edge of
the blade airfoil deviates from a right angle within a limited
range.
Inventors: |
HOFMANN; Willy Heinz;
(Baden-Ruetihof, CH) ; HUBER; Michael; (Wettingen,
CH) |
Correspondence
Address: |
BUCHANAN, INGERSOLL & ROONEY PC
POST OFFICE BOX 1404
ALEXANDRIA
VA
22313-1404
US
|
Assignee: |
ALSTOM Technology Ltd
Baden
CH
|
Family ID: |
39870460 |
Appl. No.: |
12/892555 |
Filed: |
September 28, 2010 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
PCT/EP2009/052533 |
Mar 4, 2009 |
|
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|
12892555 |
|
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Current U.S.
Class: |
416/241R |
Current CPC
Class: |
F05D 2240/122 20130101;
F05D 2240/304 20130101; F01D 9/02 20130101; F05D 2250/70 20130101;
F01D 5/141 20130101 |
Class at
Publication: |
416/241.R |
International
Class: |
F01D 5/14 20060101
F01D005/14 |
Foreign Application Data
Date |
Code |
Application Number |
Mar 28, 2008 |
CH |
00467/08 |
Claims
1. A blade for a rotating thermal machine, the blade comprising: a
blade airfoil which extends essentially in a radial direction, is
exposed to circumflow by an operating medium in a flow direction,
and is delimited by a leading edge and a trailing edge in the flow
direction, wherein the blade airfoil is structured such that an
angle which flow lines of the operating medium form with a shape of
the trailing edge of the blade airfoil deviates within a limited
range from a 90.degree.-angle.
2. The blade as claimed in claim 1, wherein the angle which the
flow lines form with the trailing edge of the blade airfoil is less
than 90.degree..
3. The blade as claimed in claim 1, wherein the angle which the
flow lines form with the trailing edge of the blade airfoil is more
than 90.degree..
4. The blade as claimed in claim 2, wherein the deviation of the
angle which the flow lines form with the trailing edge of the blade
airfoil lies within the range of between 0.degree. and
-10.degree..
5. The blade as claimed in claim 3, wherein the deviation of the
angle which the flow lines form with the trailing edge of the blade
airfoil lies within the range of between 0.degree. and
+10.degree..
6. The blade as claimed in claim 4, wherein the deviation of the
angle which the flow lines form with the trailing edge of the blade
airfoil lies within the range of between 0.degree. and -5.degree.
over a largest area of a height of the blade airfoil.
7. The blade as claimed in claim 5, wherein the deviation of the
angle which the flow lines form with the trailing edge of the blade
airfoil lies within the range of between 0.degree. and +5.degree.
over a largest area of a height of the blade airfoil.
8. The blade as claimed in claim 1, wherein the blade is a stator
blade.
9. A gas turbine comprising the stator blade as claimed in claim
8.
10. A gas turbine generator set with sequential combustion, the gas
turbine generator set comprising the gas turbine as claimed in
claim 9.
11. The blade as claimed in claim 1, wherein the blade has a
twisted geometry.
12. The blade as claimed in claim 4, wherein the blade is a stator
blade.
13. A gas turbine comprising the stator blade as claimed in claim
12.
14. The blade as claimed in claim 5, wherein the blade is a stator
blade.
15. A gas turbine comprising the stator blade as claimed in claim
14.
16. The blade as claimed in claim 6, wherein the blade is a stator
blade.
17. A gas turbine comprising the stator blade as claimed in claim
16.
18. The blade as claimed in claim 7, wherein the blade is a stator
blade.
19. A gas turbine comprising the stator blade as claimed in claim
18.
20. The blade as claimed in claim 8, wherein the stator blade has a
twisted geometry.
Description
RELATED APPLICATIONS
[0001] This application claims priority as a continuation
application under 35 U.S.C. .sctn.120 to PCT/EP2009/052533, which
was filed as an International Application on Mar. 4, 2009
designating the U.S., and which claims priority to Swiss
Application 00467/08 filed in Switzerland on Mar. 28, 2008. The
entire contents of these applications are hereby incorporated by
reference in their entireties.
FIELD
[0002] The present disclosure relates to the field of thermal
machines. More particularly, the present disclosure relates to a
blade for a rotating thermal machine.
BACKGROUND INFORMATION
[0003] Large stationary gas turbines with sequential combustion
have proven to be successful in industrial use. In these gas
turbines, two combustion chambers are arranged one behind the other
in the flow direction. A turbine associated with each combustion
chamber is in each case exposed to admission of a hot gas which is
produced in the respective combustion chamber. Gas turbines of this
type, which have been known among experts, for example, under the
type designation GT24/26, follow from the printed publication by
Joos F. et al., "Field experience with the sequential combustion
system of the GT24/26 gas turbine family", ABB Review 5/1998, p.
12-20 (1998). FIG. 1 of the Joos article is reproduced in the
present disclosure as FIG. 1. A further description of such a gas
turbine also follows from EP-B1-0 620 362.
[0004] FIG. 1 shows a known gas turbine 10 with sequential
combustion, in which a compressor 11, a first combustion chamber
14, a high-pressure (HP) turbine 15, a second combustion chamber 17
and a low-pressure (LP) turbine 18 are arranged in series along an
axis 19. In such a gas turbine 10, the compressor 11 and the two
turbines 15 (HP), 18 (LP) are part of a rotor which rotates around
the axis 19. The compressor 11 compresses the inducted air, and
this compressed air then flows into a plenum and from there flows
into the first combustion chamber 14. The first combustion chamber
14 is operated with premix burners, as described in, for example,
EP-A1-0 321 809 and EP-A2-0 704 657. The compressed air flows into
the premix burners in which the compressed air is mixed with at
least one fuel. This fuel/air mixture then flows into the first
combustion chamber 14 in which this mixture is combusted, forming a
stable flame front. The resulting hot gas is partially expanded in
the adjoining high-pressure turbine 15, performing work, and then
flows into the second combustion chamber 17 in which a further fuel
feed 16 takes place. As a result of the high temperatures which the
hot gas, which is partially expanded in the high-pressure turbine
15, still has, combustion takes place in the second combustion
chamber 17, which is based on self-ignition. The hot gas which is
reheated in the second combustion chamber 17 is then expanded in
the multistage low-pressure turbine 18 in which blade rows
including rotor blades and stator blades are arranged one behind
the other in an alternating manner.
[0005] The low-pressure turbine 18 includes a blading 29 in which a
plurality of rows of rotor blades and stator blades are arranged
one behind the other and in an alternating manner in the flow
direction. The stator blades have a blade airfoil 22 (see FIG. 2)
which extends in the radial direction between a shroud 21 (see FIG.
2) and a blade tip 23 (see FIG. 2). The shroud 21 and blade tip 23
constitute two ends of the stator blade. The above-described
construction of the stator blade also applies to the rotor blades.
The two ends 21, 23 of the stator blade delimit the throughflow
cross section of a hot gas passage in the radial direction, through
which passage a hot gas flow 30 (see FIG. 2) flows and impinges
upon the blade airfoil 22 of the blade with corresponding flow
lines, of which three flow lines are exemplarily shown in FIG. 2
and designated with reference symbol 26.
[0006] As can be seen in FIG. 1, the throughflow cross section of
the hot gas passage widens considerably in the flow direction,
conforming to the turbine 10. From this, certain consequences arise
for the geometric design of the respective blade airfoil. For
instance, the blade airfoils of stator blades in gas turbines or
steam turbines are customarily designed so that the local flow
lines of the flowing operating medium (hot gas or steam) at the
point of intersection with the trailing edge of the blade airfoil
extend approximately perpendicularly to the trailing edge. On
account of the large angle of inclination of the flow path in the
meridional plane, the trailing edge, however, cannot be oriented
completely and continuously perpendicularly to the flow lines
because this would require a sharp sweep angle and tilt, for
example, at the blade tip which, however, is not possible, on
account of the given space conditions and assembly, apart from the
fact that such a configuration, even if this were able to be
accomplished, would otherwise have fluidic disadvantages.
SUMMARY
[0007] An exemplary embodiment provides a blade for a rotating
thermal machine. The exemplary blade includes a blade airfoil which
extends essentially in a radial direction, is exposed to circumflow
by an operating medium in a flow direction, and is delimited by a
leading edge and a trailing edge in the flow direction. The blade
airfoil is structured such that an angle which flow lines of the
operating medium form with a shape of the trailing edge of the
blade airfoil deviates within a limited range from a
90.degree.-angle.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] Additional refinements, advantages and features of the
present disclosure are described in more detail below with
reference to exemplary embodiments illustrated in the drawings.
Elements which are not essential for understanding the features of
the exemplary embodiments have been omitted. The same elements
and/or similarly functioning elements are provided with the same
reference symbols. The flow direction of the media is indicated by
arrows. The reference symbols used in the drawings are summarized
in the List of Reference Symbols below. In the drawings:
[0009] FIG. 1 shows a basic construction of a known gas turbine
with sequential combustion;
[0010] FIG. 2 shows in a perspective side view a stator blade, such
as one which can be used in a gas turbine according to FIG. 1, and
also shows features of a stator blade according to an exemplary
embodiment of the present disclosure;
[0011] FIG. 3 shows the deviation of an angle, which the flow lines
form with the trailing edge of a blade which is comparable to FIG.
2, from the right angle over the height of the blade if the blade
has a fully orthogonal "stacking", according to an exemplary
embodiment of the present disclosure; and
[0012] FIG. 4 shows the deviation of an angle, which the flow lines
form with the trailing edge of the blade which is shown in FIG. 2,
from the right angle over the height of the blade, according to an
exemplary embodiment of the present disclosure.
DETAILED DESCRIPTION
[0013] Exemplary embodiments of the present disclosure provide a
blade which has a fluidically optimum body shape inside a
prespecified throughflow cross section. This exemplary arrangement
can provide a maximized efficiency.
[0014] Exemplary embodiments of the present disclosure provide an
improved shape of the blade airfoil, in which the angle which the
flow lines form with the trailing edge of the blade airfoil deviate
from a right angle within a limited range, as would ensue in the
case of a constant throughflow cross section. The angle which the
flow lines therefore form with the trailing edge of the blade
airfoil can be, for example, less than 90.degree.. In some cases,
the angle can also be more than 90.degree..
[0015] According to an exemplary embodiment of the present
disclosure, the deviation with regard to this angle which the flow
lines form with the trailing edge of the blade airfoil lies within
the range of between 0.degree. and -10.degree. or +10.degree. in
comparison to a right angle. For example, the deviation of the
angle which the flow lines form with the trailing edge of the blade
airfoil can lie within the range of between 0.degree. and
-5.degree., and/or between 0.degree. and +5.degree., over the
largest area of the height of the blade airfoil. According to an
exemplary embodiment, the deviations within the angle range are not
uniform over the entire length of the blade airfoil. In other
words, according to an exemplary embodiment, the flow lines do not
have a deviation by the same amount within specific flow sections
along the length of the blade airfoil in each case. An oscillating
deviation within the applied angle range along the entire length of
the blade airfoil is also possible.
[0016] Features of the exemplary embodiments of the present
disclosure can be transferred to so-called twisted blades.
[0017] FIG. 2 shows an example of a stator blade, which can be used
in a turbine of a gas turbine generator set, for example, in the
low-pressure turbine of a gas turbine with sequential combustion
according to FIG. 1. The stator blade 20 includes a blade airfoil
22 which is comparatively sharply curved in space. The blade
airfoil extends in the longitudinal direction (in the radial
direction in relation to the rotor of the gas turbine) between a
blade tip 23 and a shroud 21, and extends from a leading edge 27 to
a trailing edge 28 in the throughflow direction (flow direction) of
the hot gas flow 30. Between the two edges 27 and 28, the blade
airfoil 22 is outwardly delimited by a suction side 31 and a
pressure side opposite to the suction side 31.
[0018] The hot gas flow 30 flows from the leading edge 27 to the
trailing edge 28 along the blade airfoil 22 in flow lines 26. Three
such flow lines 26 are exemplarily shown in FIG. 2. The flow lines
26 form an angle .alpha. at their point of intersection with the
trailing edge 28 in each case. The angle .alpha. varies in the
radial direction and thus has a dependency upon the height h of the
blade airfoil 22.
[0019] If this angle .alpha. is equal to 90.degree. over the entire
height of the blade airfoil 22, this would correspond to a "fully
orthogonal stacking" of the blades. Therefore, for the deviation
.alpha.-90.degree. from the right angle, the value 0 would
continuously result, as is shown in the diagram of FIG. 3, in which
the function .alpha.-90.degree. (h) is displayed.
[0020] According to an exemplary embodiment of the present
disclosure, this fully orthogonal stacking is replaced by a less
strict "relaxed orthogonal stacking", in which the angle .alpha.
indeed stays close to a right angle but can deviate from this
within a limited range. The diagram corresponding to FIG. 3 for
such a "relaxed" orthogonal stacking is reproduced in FIG. 4. The
deviation .alpha.-90.degree. of the angle .alpha. from the right
angle lies within the negative range (<0) and overall is limited
to an angle range of between 0 and -10.degree. in the currently
shown example. At the lower (rotor-side) end of the blade airfoil
22 (small h), the deviation begins with a maximum value of almost
-10.degree., then returns to zero within a very short distance, and
in the currently shown example remains below -5.degree. over the
largest area of the height. As a result, with this flexible
maintained deviation, special flow conditions in the sense of
maximizing the efficiency can be gained.
[0021] According to this exemplary arrangement, a limited deviation
of this type from the strict orthogonal stacking at the same time
can meet the fluidic, constructional and space-dependent demands on
the blade inside a varying throughflow cross section, including
maximizing the efficiency.
[0022] It will be appreciated by those skilled in the art that the
present invention can be embodied in other specific forms without
departing from the spirit or essential characteristics thereof. The
presently disclosed embodiments are therefore considered in all
respects to be illustrative and not restricted. The scope of the
invention is indicated by the appended claims rather than the
foregoing description and all changes that come within the meaning
and range and equivalence thereof are intended to be embraced
therein.
LIST OF DESIGNATIONS
[0023] 10 Gas turbine [0024] 11 Compressor [0025] 12, 16 Fuel feed
[0026] 13 EV-burner [0027] 14, 17 Combustion chamber [0028] 15
High-pressure turbine [0029] 18 Low-pressure turbine [0030] 19 Axis
[0031] 20 Stator blade [0032] 21 Shroud [0033] 22 Blade airfoil
[0034] 23 Blade tip [0035] 24, 25 Fastening element (hook-like)
[0036] 26 Flow line [0037] 27 Leading edge [0038] 28 Trailing edge
[0039] 29 Blading (low-pressure turbine) [0040] 30 Hot gas flow
[0041] 31 Suction side [0042] .alpha. Angle
* * * * *