CA1125660A - Cooled vane structure for a combustion turbine engine - Google Patents
Cooled vane structure for a combustion turbine engineInfo
- Publication number
- CA1125660A CA1125660A CA353,911A CA353911A CA1125660A CA 1125660 A CA1125660 A CA 1125660A CA 353911 A CA353911 A CA 353911A CA 1125660 A CA1125660 A CA 1125660A
- Authority
- CA
- Canada
- Prior art keywords
- airfoil
- vane
- cooling air
- cavity
- sidewalls
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
8 48,590 ABSTRACT OF THE DISCLOSURE
A cooled vane for a combustion turbine is shown having a generally hollow airfoil for receiving cooling air and with the sidewalls of the airfoil tapered in thickness from thick-to-thin in the radial direction from the vane support into the generally free end.
A cooled vane for a combustion turbine is shown having a generally hollow airfoil for receiving cooling air and with the sidewalls of the airfoil tapered in thickness from thick-to-thin in the radial direction from the vane support into the generally free end.
Description
COOLED VANE STRUCTURE FOR
COMBIJSTION TURBINE ENGINE
BACKGROUND OF THE INVENTION
Field of the Invention:
~ This invention relates to the stator vane struc-`~ ture for generally the first stage of a multi-stage axial flow gas turbine.
. Description_of the_Prior Art:
.
: U.S. Patent No. 3,689,174, of common assignee to the present invention, describes a single vane structure for the first stage of a multi-stage gas turbine engine.
As is noted, this prior art vane is generally hollow, permitting cooling air to enter its core to cool the vane walls. It is apparent that the airfoil portion of the first row of vanes are the first to receive the hot motive `~ - gases discharged from the combustion chamber of the engine and thus are directly subjected to the high temperatures and also the the high velocity of such gases such that, unless the vane is cooled, these temperatures and stresses will cause premature failure of the vane structure.
However, it was found that the general hollow ~` 20 construction, such as in the above-identified patent, did not provide the desired specific zone cooling (i.e. some - areas or zones of the vane walls, such as the nose or :~ leading edge require more cooling than others) and thus, to obtain sufficiently cool areas in such zones, an excess ~, 25 of cooling air was required.
Another known cooled vane structure provided :, . . -. - ~
- . . . .
radially extending individual cooling holes through the airfoil section, permitting more selec-tive distribution of the cooling air (i.e. more cooling holes could be provided in those areas subjected most directly to the extreme temperatures and which therefore required the most cool-ing). However, these holes, in the casting process, were difficult to control and resulted in a high scrap rate and thus such vane configuration became too expensive to manufacture.
~
The vane configuration of the present invention reduces the operating metal temperature, in order to im-prove vane life, by reducing the airfoil wall thickness and tapering the wall thickness such that it increases with the flow path radius (which is also in the direction towards the vane support) just sufficient to support the stress in the airfoil which likewise increases with each increase in flow path radius.
Thus, from the direction of the greatest radius which devines the outer shroud portion of the vane to the innermost radius at the inner shroud portion, the airfoil wall thickness decreases; however, the heat removal capa-city of the incoming cooling alr entering from a chamber adjacent the outer shroud is greatest at its initial entry which conforms to the greatest wall thickness and which, in turn, requires the most cooling. As the cooling air progresses radially i~wardly and becomes heated, the wall thickness decreases and thus requires less heat removal.
.~ Such tapered wall configuration reduces the amount of material required to fabricate the vane and also reduces ` the mass of material from which heat must be removed which, in turn, minimizes the amount of cooling air re-quired.
Further, the distribution of the vane cooling air is controlled and its effectiveness is increased by an insert disposed within the vane cavity, for initial re-ceipt of the cooling, air and having ~a hole pattern for discharging the cooling air in impin~ement jets against ."
l~'Æ5~60 3 4~" 590 the walls of ~he airfoil portion. After cooling the vane walls, the cooling air is exhaus-ted from the vane ca~it~
through traîling edge slots machined in the trailing edge of -the airfoil, BRIEF DESCRIPTION OF THE DRAI~IN~S
Figure 1 ls an isometric view of ~he vane of the present invention;
Figo 2 is a top plan v~ew of Fig~ l;
Fig. 3 is a cross-sectional view ge~erally along the lines III-III of Fig~ 2; a~d Fig. 4 is a cross-sectional view gener~lly along line IV-IV of Figo 2~
Referring ~o ~he Figures, khe vane 10 of the present invention is seen to include an outer shroud 12 and an inner shroud 1~. The upper surface 1~ of the outer shroud 12 provides structure for mounting the vane within the combustion turbine similar to the mounting arrangement more fully described in the co-owned patent previously identified and more particularly described in Canadian co~only owned patent application Serial ~o. 353,910 ~iled o~ Jlme 12~ 19~0.
The mounting s~ructure in vane 10 includes opposed facing rail members 1~ 20 having tongue portions 229 2~ ~or tongue-in-groove engagement with a slot in the mounting blocks (no~ shown~ in the engine. A ~enerally ellipti~al shoulder 26 elevated ~rom the general plane o~
~he upper sur~ace 16 deflnes a surface for indexed mount-ing ~as b~ brazing) of a cover member 2~ to provîde a chamber 29 for reeeiving the vane cooling air.
me air~oil portion 30 o~ the ~ane 10 extends between ~he opposed shrouds 12, 14 and de~ines a nose or ~eading edge portion 32 and sidewalls 34 (pr~ssure side) 36 (suo~lon side) converging to a trailing edge 3~ The ; 35 airfoil por~ion 30 is generally hollow having an airfoil shaped opening 40 extending radially therethrough and through the upper shroud within the boundary of the ellip-~ tical shoulder 26 to place the cavity 40 in cooling air .
.
' :
.
- .
6i~
flow communication with the cooling air delivered through the cover member 28 to the chamber 29.
The upper periphery of the cavi~y 40 provides a raised lip 42 and an airfoil-shaped insert 44 having a top peripheral flange 46 is positioned in spaced relationship within ~he opening ~lO with the ~lange welded to the lip 42. The insert 44 provides a plurality of openings 43 therethrough in a preselected pattern to deliver impinge-ment jets of cooling air to the inner surface of the cavity 40 to cool the walls of the airfoil portion and provide the desired cooling air distribution. Trailing edge slots 48 lead from the cavity 40 to the exterior of the vane 10 through the trailing edge to finally exhaust the cooling air to the motive gas path o~ the engine.
15Referring now more particularly to Fig. 3 it is seen that the sidewalls (i.e. the pressure sidewall 34 and the suction sidewall 36) have a tapered thickness, com-mencing at the outer shroud 12 in a relatively thick cross section and gradually becoming increasingly less thick in the direction of the inner shroud 14. This tapered side-wall thickness (i.e. it is noted in Fig. 4 that the thick-; ness of the nose and the trail:ing edge remain generally constant) provides the cross-sectional area suficient to support the increased stress in the vane as each radially outer area must support proportionally greater stress, ` such as bending stresses, so that gradually more area is req~lired. The tapered structure maintains a minimum wall thickness for such support. This, in turn, results in less vane material to be cooled so that the minimum amount of cooling air must be supplied to the insert. Further, since the vane is supported from the same end from which the cooling air is delivered (i.e. the upper shroud 1~
the coolest cooling air is delivered through the insert to the thickest wall area of the airfoil portion 30 so that the heat removal by the impingement of air thereon is the most efficient in the area having the greatest material from which the greatest amount of heat is removed. Also, ~: the hole pattern in the insert is such as to deliver ' :' - - .
..
.
~ 2~6~
cooling air to accommodate the heat removal from the vane in general conformity to the temperature profile across t-he vane. ~ C~ S~
Thus, a cooled vane is shown for thc c~bin~ti~;
turbine engine, having a hollow airfoil portion with sidewalls have a thickness tapering from their thickest area adjacent the vane support to their thinnest area adjacent the generally free end. Cooling air also enters the internal cavity of the airfoil portion from the end thereof having the thickest sidewalls. This ultimately minimizes the material in the airfoil por-tion of the vane which, in turn, reduces the amount of cooling air flow : through the vane necessary to cool the airfoil walls to an acceptable temperature level.
- . 'I . ' ' ' ~, .
.
: ,, ' ' : ' . ~ . .
:. . , -. : .
. ' - ' : '
COMBIJSTION TURBINE ENGINE
BACKGROUND OF THE INVENTION
Field of the Invention:
~ This invention relates to the stator vane struc-`~ ture for generally the first stage of a multi-stage axial flow gas turbine.
. Description_of the_Prior Art:
.
: U.S. Patent No. 3,689,174, of common assignee to the present invention, describes a single vane structure for the first stage of a multi-stage gas turbine engine.
As is noted, this prior art vane is generally hollow, permitting cooling air to enter its core to cool the vane walls. It is apparent that the airfoil portion of the first row of vanes are the first to receive the hot motive `~ - gases discharged from the combustion chamber of the engine and thus are directly subjected to the high temperatures and also the the high velocity of such gases such that, unless the vane is cooled, these temperatures and stresses will cause premature failure of the vane structure.
However, it was found that the general hollow ~` 20 construction, such as in the above-identified patent, did not provide the desired specific zone cooling (i.e. some - areas or zones of the vane walls, such as the nose or :~ leading edge require more cooling than others) and thus, to obtain sufficiently cool areas in such zones, an excess ~, 25 of cooling air was required.
Another known cooled vane structure provided :, . . -. - ~
- . . . .
radially extending individual cooling holes through the airfoil section, permitting more selec-tive distribution of the cooling air (i.e. more cooling holes could be provided in those areas subjected most directly to the extreme temperatures and which therefore required the most cool-ing). However, these holes, in the casting process, were difficult to control and resulted in a high scrap rate and thus such vane configuration became too expensive to manufacture.
~
The vane configuration of the present invention reduces the operating metal temperature, in order to im-prove vane life, by reducing the airfoil wall thickness and tapering the wall thickness such that it increases with the flow path radius (which is also in the direction towards the vane support) just sufficient to support the stress in the airfoil which likewise increases with each increase in flow path radius.
Thus, from the direction of the greatest radius which devines the outer shroud portion of the vane to the innermost radius at the inner shroud portion, the airfoil wall thickness decreases; however, the heat removal capa-city of the incoming cooling alr entering from a chamber adjacent the outer shroud is greatest at its initial entry which conforms to the greatest wall thickness and which, in turn, requires the most cooling. As the cooling air progresses radially i~wardly and becomes heated, the wall thickness decreases and thus requires less heat removal.
.~ Such tapered wall configuration reduces the amount of material required to fabricate the vane and also reduces ` the mass of material from which heat must be removed which, in turn, minimizes the amount of cooling air re-quired.
Further, the distribution of the vane cooling air is controlled and its effectiveness is increased by an insert disposed within the vane cavity, for initial re-ceipt of the cooling, air and having ~a hole pattern for discharging the cooling air in impin~ement jets against ."
l~'Æ5~60 3 4~" 590 the walls of ~he airfoil portion. After cooling the vane walls, the cooling air is exhaus-ted from the vane ca~it~
through traîling edge slots machined in the trailing edge of -the airfoil, BRIEF DESCRIPTION OF THE DRAI~IN~S
Figure 1 ls an isometric view of ~he vane of the present invention;
Figo 2 is a top plan v~ew of Fig~ l;
Fig. 3 is a cross-sectional view ge~erally along the lines III-III of Fig~ 2; a~d Fig. 4 is a cross-sectional view gener~lly along line IV-IV of Figo 2~
Referring ~o ~he Figures, khe vane 10 of the present invention is seen to include an outer shroud 12 and an inner shroud 1~. The upper surface 1~ of the outer shroud 12 provides structure for mounting the vane within the combustion turbine similar to the mounting arrangement more fully described in the co-owned patent previously identified and more particularly described in Canadian co~only owned patent application Serial ~o. 353,910 ~iled o~ Jlme 12~ 19~0.
The mounting s~ructure in vane 10 includes opposed facing rail members 1~ 20 having tongue portions 229 2~ ~or tongue-in-groove engagement with a slot in the mounting blocks (no~ shown~ in the engine. A ~enerally ellipti~al shoulder 26 elevated ~rom the general plane o~
~he upper sur~ace 16 deflnes a surface for indexed mount-ing ~as b~ brazing) of a cover member 2~ to provîde a chamber 29 for reeeiving the vane cooling air.
me air~oil portion 30 o~ the ~ane 10 extends between ~he opposed shrouds 12, 14 and de~ines a nose or ~eading edge portion 32 and sidewalls 34 (pr~ssure side) 36 (suo~lon side) converging to a trailing edge 3~ The ; 35 airfoil por~ion 30 is generally hollow having an airfoil shaped opening 40 extending radially therethrough and through the upper shroud within the boundary of the ellip-~ tical shoulder 26 to place the cavity 40 in cooling air .
.
' :
.
- .
6i~
flow communication with the cooling air delivered through the cover member 28 to the chamber 29.
The upper periphery of the cavi~y 40 provides a raised lip 42 and an airfoil-shaped insert 44 having a top peripheral flange 46 is positioned in spaced relationship within ~he opening ~lO with the ~lange welded to the lip 42. The insert 44 provides a plurality of openings 43 therethrough in a preselected pattern to deliver impinge-ment jets of cooling air to the inner surface of the cavity 40 to cool the walls of the airfoil portion and provide the desired cooling air distribution. Trailing edge slots 48 lead from the cavity 40 to the exterior of the vane 10 through the trailing edge to finally exhaust the cooling air to the motive gas path o~ the engine.
15Referring now more particularly to Fig. 3 it is seen that the sidewalls (i.e. the pressure sidewall 34 and the suction sidewall 36) have a tapered thickness, com-mencing at the outer shroud 12 in a relatively thick cross section and gradually becoming increasingly less thick in the direction of the inner shroud 14. This tapered side-wall thickness (i.e. it is noted in Fig. 4 that the thick-; ness of the nose and the trail:ing edge remain generally constant) provides the cross-sectional area suficient to support the increased stress in the vane as each radially outer area must support proportionally greater stress, ` such as bending stresses, so that gradually more area is req~lired. The tapered structure maintains a minimum wall thickness for such support. This, in turn, results in less vane material to be cooled so that the minimum amount of cooling air must be supplied to the insert. Further, since the vane is supported from the same end from which the cooling air is delivered (i.e. the upper shroud 1~
the coolest cooling air is delivered through the insert to the thickest wall area of the airfoil portion 30 so that the heat removal by the impingement of air thereon is the most efficient in the area having the greatest material from which the greatest amount of heat is removed. Also, ~: the hole pattern in the insert is such as to deliver ' :' - - .
..
.
~ 2~6~
cooling air to accommodate the heat removal from the vane in general conformity to the temperature profile across t-he vane. ~ C~ S~
Thus, a cooled vane is shown for thc c~bin~ti~;
turbine engine, having a hollow airfoil portion with sidewalls have a thickness tapering from their thickest area adjacent the vane support to their thinnest area adjacent the generally free end. Cooling air also enters the internal cavity of the airfoil portion from the end thereof having the thickest sidewalls. This ultimately minimizes the material in the airfoil por-tion of the vane which, in turn, reduces the amount of cooling air flow : through the vane necessary to cool the airfoil walls to an acceptable temperature level.
- . 'I . ' ' ' ~, .
.
: ,, ' ' : ' . ~ . .
:. . , -. : .
. ' - ' : '
Claims
1. A cooled vane assembly for a combustion turbine engine comprising:
an inner shroud;
an outer shroud;
an airfoil portion supported by said outer shroud and having a leading edge and a trailling edge and opposed pressure and suction sides and further having an internal cavity, extending between said edges and said sides, said pressure and suction sidewalls having a varying thickness radially along said pressure and suction sides which is tapered from thick-to-thin in a radially inward direction from said outer shroud to said inner shroud;
means in said outer shroud defining a cooling air inlet to said cavity;
an airfoil insert supported in said airfoil cavity between said inner and outer shrouds and having a wall member surrounding an inner coolant cavity and extending along said airfoil sidewalls in spaced relation therewith and between said airfoil edges, said insert having a plurality of holes to supply coolant air to the interwall space between it and said airfoil sidewalls and edges; and means for exhausting cooling air from the airfoil interwall space.
an inner shroud;
an outer shroud;
an airfoil portion supported by said outer shroud and having a leading edge and a trailling edge and opposed pressure and suction sides and further having an internal cavity, extending between said edges and said sides, said pressure and suction sidewalls having a varying thickness radially along said pressure and suction sides which is tapered from thick-to-thin in a radially inward direction from said outer shroud to said inner shroud;
means in said outer shroud defining a cooling air inlet to said cavity;
an airfoil insert supported in said airfoil cavity between said inner and outer shrouds and having a wall member surrounding an inner coolant cavity and extending along said airfoil sidewalls in spaced relation therewith and between said airfoil edges, said insert having a plurality of holes to supply coolant air to the interwall space between it and said airfoil sidewalls and edges; and means for exhausting cooling air from the airfoil interwall space.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US5362979A | 1979-06-29 | 1979-06-29 | |
| US053,629 | 1979-06-29 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1125660A true CA1125660A (en) | 1982-06-15 |
Family
ID=21985545
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA353,911A Expired CA1125660A (en) | 1979-06-29 | 1980-06-12 | Cooled vane structure for a combustion turbine engine |
Country Status (2)
| Country | Link |
|---|---|
| JP (2) | JPS569623A (en) |
| CA (1) | CA1125660A (en) |
Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4511306A (en) * | 1982-02-02 | 1985-04-16 | Westinghouse Electric Corp. | Combustion turbine single airfoil stator vane structure |
| JP2602123B2 (en) * | 1990-10-18 | 1997-04-23 | 株式会社堀場製作所 | Car driving robot on chassis dynamometer |
| JP2566448Y2 (en) * | 1991-09-10 | 1998-03-25 | 株式会社堀場製作所 | Car driving robot on chassis dynamometer |
| JP2537226Y2 (en) * | 1991-09-10 | 1997-05-28 | 株式会社堀場製作所 | Car driving robot on chassis dynamometer |
| FR2899271B1 (en) * | 2006-03-29 | 2008-05-30 | Snecma Sa | DUSTBOARD AND COOLING SHIELD ASSEMBLY, TURBOMACHINE DISPENSER COMPRISING THE ASSEMBLY, TURBOMACHINE, METHOD OF ASSEMBLING AND REPAIRING THE ASSEMBLY |
| CN119744324A (en) * | 2022-09-05 | 2025-04-01 | 三菱重工业株式会社 | Cooling fluid guide for gas turbine and gas turbine |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3275294A (en) * | 1963-11-14 | 1966-09-27 | Westinghouse Electric Corp | Elastic fluid apparatus |
| GB1564608A (en) * | 1975-12-20 | 1980-04-10 | Rolls Royce | Means for cooling a surface by the impingement of a cooling fluid |
-
1980
- 1980-06-12 CA CA353,911A patent/CA1125660A/en not_active Expired
- 1980-06-27 JP JP8673480A patent/JPS569623A/en active Pending
-
1985
- 1985-07-02 JP JP10016185U patent/JPS6117401U/en active Pending
Also Published As
| Publication number | Publication date |
|---|---|
| JPS6117401U (en) | 1986-01-31 |
| JPS569623A (en) | 1981-01-31 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| MKEX | Expiry |