US20110038733A1

US20110038733A1 - Blade for a rotating thermal machine

Info

Publication number: US20110038733A1
Application number: US12/892,555
Authority: US
Inventors: Willy Heinz Hofmann; Michael Huber
Original assignee: Alstom Technology AG
Current assignee: General Electric Technology GmbH
Priority date: 2008-03-28
Filing date: 2010-09-28
Publication date: 2011-02-17
Also published as: WO2009118234A1; EP2268900A1

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.

Description

RELATED APPLICATIONS

This application claims priority as a continuation application under 35 U.S.C. §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

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

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.
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.
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.
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

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°-angle.

BRIEF DESCRIPTION OF THE DRAWINGS

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:

FIG. 1 shows a basic construction of a known gas turbine with sequential combustion;

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;

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

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

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.
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°. In some cases, the angle can also be more than 90°.
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° and −10° or +10° 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° and −5°, and/or between 0° and +5°, 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.
Features of the exemplary embodiments of the present disclosure can be transferred to so-called twisted blades.
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.
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 α at their point of intersection with the trailing edge 28 in each case. The angle α varies in the radial direction and thus has a dependency upon the height h of the blade airfoil 22.
If this angle α is equal to 90° over the entire height of the blade airfoil 22, this would correspond to a “fully orthogonal stacking” of the blades. Therefore, for the deviation α-90° from the right angle, the value 0 would continuously result, as is shown in the diagram of FIG. 3, in which the function α-90° (h) is displayed.
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 α 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 α-90° of the angle α from the right angle lies within the negative range (<0) and overall is limited to an angle range of between 0 and −10° 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°, then returns to zero within a very short distance, and in the currently shown example remains below −5° 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.
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.
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

10 Gas turbine
11 Compressor
12, 16 Fuel feed
13 EV-burner
14, 17 Combustion chamber
15 High-pressure turbine
18 Low-pressure turbine
19 Axis
20 Stator blade
21 Shroud
22 Blade airfoil
23 Blade tip
24, 25 Fastening element (hook-like)
26 Flow line
27 Leading edge
28 Trailing edge
29 Blading (low-pressure turbine)
30 Hot gas flow
31 Suction side
α Angle

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°-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°.

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°.

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° and −10°.

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° and +10°.

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° and −5° 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° and +5° 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.