CA1183695A - Efficiently cooled transition duct for a large plant combustion turbine - Google Patents
Efficiently cooled transition duct for a large plant combustion turbineInfo
- Publication number
- CA1183695A CA1183695A CA000388209A CA388209A CA1183695A CA 1183695 A CA1183695 A CA 1183695A CA 000388209 A CA000388209 A CA 000388209A CA 388209 A CA388209 A CA 388209A CA 1183695 A CA1183695 A CA 1183695A
- Authority
- CA
- Canada
- Prior art keywords
- duct
- skin
- tubular
- shell
- coolant
- 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
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
ABSTRACT OF THE DISCLOSURE
A transition duct for a large plant combustion turbine comprises an elongated outer tubular wall member arranged to be coupled with a combustion chamber at its upstream end and to be supported at its downstream end to direct the hot driving gases through the annular turbine blade space. The tubular wall has at least a portion thereof structured for internal wall cooling including an outer tubular shell member and a skin member having a continuous inner tubular surface and disposed peripherally about and against an inward facing tubular surface of the shell member. Coolant channels are arranged peripherally about and in the skin member to face the shell member and to direct coolant generally longitudinally of the duct portion. The shell member has entry means for directing compressor discharge coolant air from the duct exterior to and through the skin coolant channels. The coolant is discharged from the channels to the duct gas flow from at least one location longitudinally spaced from the entry means.
The entry means comprises at least one substan-tially transverse coolant channel disposed about said tubular shell member to face inwardly toward the tubular skin member. A plurality of openings extend from the shell coolant channel through the shell wall to the duct exterior are disposed about the shell member.
The channels extend longitudinally between the upstream and downstream ends of said internally cooled turbine wall portion, and the discharge means includes openings from the skin channels through the downstream end of the skin member.
A transition duct for a large plant combustion turbine comprises an elongated outer tubular wall member arranged to be coupled with a combustion chamber at its upstream end and to be supported at its downstream end to direct the hot driving gases through the annular turbine blade space. The tubular wall has at least a portion thereof structured for internal wall cooling including an outer tubular shell member and a skin member having a continuous inner tubular surface and disposed peripherally about and against an inward facing tubular surface of the shell member. Coolant channels are arranged peripherally about and in the skin member to face the shell member and to direct coolant generally longitudinally of the duct portion. The shell member has entry means for directing compressor discharge coolant air from the duct exterior to and through the skin coolant channels. The coolant is discharged from the channels to the duct gas flow from at least one location longitudinally spaced from the entry means.
The entry means comprises at least one substan-tially transverse coolant channel disposed about said tubular shell member to face inwardly toward the tubular skin member. A plurality of openings extend from the shell coolant channel through the shell wall to the duct exterior are disposed about the shell member.
The channels extend longitudinally between the upstream and downstream ends of said internally cooled turbine wall portion, and the discharge means includes openings from the skin channels through the downstream end of the skin member.
Description
3~
1 4g,g31 EFFICIENTLY COOLED TRANSITION DUCT FOR A
LARGE PLANT COMBUSTION TURBINE
BACKGROUND OF THE INVENTION
The present invention relates to large combus-tion turbines for industrial process and electric power generation usage and more particularly to transition ducts employed therein to direct the hot driving gases from the turbine combustors to the turbine blades.
As distinguished from aircraft engines in which the combustors are compactly arranged about the engine axis, the large plant combustion turbine structural design requires that combustors be peripherally arranged about the turbine longitudinal axis at a comparatively outward radial location. Typically, each combustor is couplad to a transition duct which is structured at its inlet end with a circular cross-section like that of the combustor and at its downstream outlet end with a cross-section corresponding to a section of the annular space through which the turbine blades rotate. Further, the transition duct normally extend,s radially inwardl~ along its length so that the combustor yases are directed from the radially outward combustor location to the more inward radial location of the downstream turbine blades.
Even with the use of improv ~ stainless steel or nickel based alloys such as Hastalloy or Inco 617~E9it has ~~ become increasingly difficult to provide adequate cooling o~ transition ducts to assure long duct life and, to some extent, long blade and vane life. Thus, gas operating ~ 48,931 temperatures have been increasing through the years to obtain higher power generation more efficiently. At present, the temperature OI operating gases is typically 2300F while the temperature operating limit for special metals used for transition ducts is about 1500~F. It is expected that turbine gas operating temperatures will be increased even further in the future.
Cooling of transition ducts is presently pro vided by compressor discharge air which is circulated about the outer duct surfaces prior to intake into the combustor baskets where it supports combustion of the liquid or gaseous fuel. Such external duct cooling has generally been adequate. However, it is becoming less so with increasing turbine gas operating temperatures for the reason already indicated.
Moreover, external duct cooling has in certain respects been inadequate in the past as well. Thus, even at current and lower turbine gas operating temperatures, some transition ducts have failed in operation because of 2C thermal stress fatique especially at bends. Duct bends and similar structural features would have even more reduced creep life with increased operating temperatures.
Further, failures have occurred in the duct portion having an annulus section at the duct downstream end since it is not as structurally effective as the duct portion having the circular cross-section at the duct upstream end at withstanding thermal stresses and pressure differentials from the inside of the duct to the outside of the duct.
To provide additional duct wall cooling, coolant air can be introduced along the inside surfaces of the duct walls to provide internal boundary layer cooling.
However, this approach distorts the temperature profile across the gas stream which causes a shortening of turbine blade and vane life. Further, such use of coolant air reduces the machine efficiency.
`~ 3~
3 ~8,931 ~UMMARY OF THE INVENTION
A transition duct for a large plant combustion turbine comprises an elongated outer tubular wall arranged to be coupled with a combustion chamber at its upstream end and to be supported at its downstream end to direct the hot driving gases through the annular turbine blade space. At least a portion of the tubular wall is struc-tured for internal wall cooling by including an outer tubular shell member and a skin member preferably having a continuous inner tubular surface and an outer surface secured peripherally about and against an inward facing tubular surface of the shell member. Coolant channel means are arranged peripherally about and in the skin member to face the shell member and to direct coolant generally longitudinally of the duct portion between the shell and skin members. The shell member is provided with entry means for directing compressor discharge coolant air from the duct exterior to and through the skin coolant channel means, and means are provided for discharging the coolant from the channel means to the duct internal gas flow from at least one location longitudinally spaced from the entry means.
BRIEF DESCRIPTION OF THE DRAWINGS
F:igures l shows a longitudinal section of a combustion turbine in which a transition duct is arranged to be cooled in accordance with the principles of the invention;
Figure 2 shows a~ elevational view of the tran~
sition duct along a longitudinal section thereof;
Figure 3 shows an upstream end view of an entry portion of the duct;
Figure 4 shows a downstream end view of the discharge or mouth portion of the duct;
Figure 5 shows an enlarged view of area C of the downstream ~uct mouth portion shown in Figure 4;
Figures 6 and 7 show enlarged views of respec-tive areas A and B of the longitudinal duct section of Figure ~;
336~
4 48,931 Figure 8 shows a portion of the duct cross-section along section line VIII-VIII of Figure 6;
Figure 9 graphically illustrates the metal tem-perature performance of a duct constructed in accordance with the invention;
Figure 10 shows an enlarged cross-section through a downstream portion of another duct embodiment shown in Figure 12;
Figure 11 shows section XI-XI o~ Figure 10;
Figure 12 shows another embodiment comprising a duct employing transpirating skin structure to provide enhanced cooling;
Figure 13 shows an enlarged view of area D in Figure 12;
Figure 14 shows an enlarged view of area E in Figure 12; and Figure 15 shows a top view of the enlarged duct area of Figure 14.
DESCRIPTION OF THE PREFERRED EMBODIMENT
~ore particularly, there is shown in Figure 1 a large plant combustion turbine 10 having a plurality of generally cylindrical combustors 12. Fuel is supplied to the combustors through a nozzle structure 14 and air is supplied to the combustors 12 by a compressor 16 through air flow space 18 within a combustion turbine casing 20.
Hot gaseous products of com~ustion are directed from each combustor 12 through an upstream end of a tran-sition duct 22 to a downstream duct end where it is dis-charged into the annular space through which turbine blades rotate under the driving ~orce o~ the expanding gases. In this case, three blade stages (only two stages
1 4g,g31 EFFICIENTLY COOLED TRANSITION DUCT FOR A
LARGE PLANT COMBUSTION TURBINE
BACKGROUND OF THE INVENTION
The present invention relates to large combus-tion turbines for industrial process and electric power generation usage and more particularly to transition ducts employed therein to direct the hot driving gases from the turbine combustors to the turbine blades.
As distinguished from aircraft engines in which the combustors are compactly arranged about the engine axis, the large plant combustion turbine structural design requires that combustors be peripherally arranged about the turbine longitudinal axis at a comparatively outward radial location. Typically, each combustor is couplad to a transition duct which is structured at its inlet end with a circular cross-section like that of the combustor and at its downstream outlet end with a cross-section corresponding to a section of the annular space through which the turbine blades rotate. Further, the transition duct normally extend,s radially inwardl~ along its length so that the combustor yases are directed from the radially outward combustor location to the more inward radial location of the downstream turbine blades.
Even with the use of improv ~ stainless steel or nickel based alloys such as Hastalloy or Inco 617~E9it has ~~ become increasingly difficult to provide adequate cooling o~ transition ducts to assure long duct life and, to some extent, long blade and vane life. Thus, gas operating ~ 48,931 temperatures have been increasing through the years to obtain higher power generation more efficiently. At present, the temperature OI operating gases is typically 2300F while the temperature operating limit for special metals used for transition ducts is about 1500~F. It is expected that turbine gas operating temperatures will be increased even further in the future.
Cooling of transition ducts is presently pro vided by compressor discharge air which is circulated about the outer duct surfaces prior to intake into the combustor baskets where it supports combustion of the liquid or gaseous fuel. Such external duct cooling has generally been adequate. However, it is becoming less so with increasing turbine gas operating temperatures for the reason already indicated.
Moreover, external duct cooling has in certain respects been inadequate in the past as well. Thus, even at current and lower turbine gas operating temperatures, some transition ducts have failed in operation because of 2C thermal stress fatique especially at bends. Duct bends and similar structural features would have even more reduced creep life with increased operating temperatures.
Further, failures have occurred in the duct portion having an annulus section at the duct downstream end since it is not as structurally effective as the duct portion having the circular cross-section at the duct upstream end at withstanding thermal stresses and pressure differentials from the inside of the duct to the outside of the duct.
To provide additional duct wall cooling, coolant air can be introduced along the inside surfaces of the duct walls to provide internal boundary layer cooling.
However, this approach distorts the temperature profile across the gas stream which causes a shortening of turbine blade and vane life. Further, such use of coolant air reduces the machine efficiency.
`~ 3~
3 ~8,931 ~UMMARY OF THE INVENTION
A transition duct for a large plant combustion turbine comprises an elongated outer tubular wall arranged to be coupled with a combustion chamber at its upstream end and to be supported at its downstream end to direct the hot driving gases through the annular turbine blade space. At least a portion of the tubular wall is struc-tured for internal wall cooling by including an outer tubular shell member and a skin member preferably having a continuous inner tubular surface and an outer surface secured peripherally about and against an inward facing tubular surface of the shell member. Coolant channel means are arranged peripherally about and in the skin member to face the shell member and to direct coolant generally longitudinally of the duct portion between the shell and skin members. The shell member is provided with entry means for directing compressor discharge coolant air from the duct exterior to and through the skin coolant channel means, and means are provided for discharging the coolant from the channel means to the duct internal gas flow from at least one location longitudinally spaced from the entry means.
BRIEF DESCRIPTION OF THE DRAWINGS
F:igures l shows a longitudinal section of a combustion turbine in which a transition duct is arranged to be cooled in accordance with the principles of the invention;
Figure 2 shows a~ elevational view of the tran~
sition duct along a longitudinal section thereof;
Figure 3 shows an upstream end view of an entry portion of the duct;
Figure 4 shows a downstream end view of the discharge or mouth portion of the duct;
Figure 5 shows an enlarged view of area C of the downstream ~uct mouth portion shown in Figure 4;
Figures 6 and 7 show enlarged views of respec-tive areas A and B of the longitudinal duct section of Figure ~;
336~
4 48,931 Figure 8 shows a portion of the duct cross-section along section line VIII-VIII of Figure 6;
Figure 9 graphically illustrates the metal tem-perature performance of a duct constructed in accordance with the invention;
Figure 10 shows an enlarged cross-section through a downstream portion of another duct embodiment shown in Figure 12;
Figure 11 shows section XI-XI o~ Figure 10;
Figure 12 shows another embodiment comprising a duct employing transpirating skin structure to provide enhanced cooling;
Figure 13 shows an enlarged view of area D in Figure 12;
Figure 14 shows an enlarged view of area E in Figure 12; and Figure 15 shows a top view of the enlarged duct area of Figure 14.
DESCRIPTION OF THE PREFERRED EMBODIMENT
~ore particularly, there is shown in Figure 1 a large plant combustion turbine 10 having a plurality of generally cylindrical combustors 12. Fuel is supplied to the combustors through a nozzle structure 14 and air is supplied to the combustors 12 by a compressor 16 through air flow space 18 within a combustion turbine casing 20.
Hot gaseous products of com~ustion are directed from each combustor 12 through an upstream end of a tran-sition duct 22 to a downstream duct end where it is dis-charged into the annular space through which turbine blades rotate under the driving ~orce o~ the expanding gases. In this case, three blade stages (only two stages
2~, 26 are shown) are pro~ided along with corresponding stationary vanes 30, 32.
Compressor discharge air provides external cooling ~or the tu~ular transition duct wall as it flows therearound as indicated by the illustrated ~low paths in Figure 1.
~ ~ ~3i~;i9~
48,931 ~ o provide high turbine operating efficiency and extended transition duct and blade and vane operating life, a transition duct is structured with enhanced wall cooling in accordance with the principles of the inven-tion. The preferred structure for the transition duct 22is shown in Figures 2-8.
The duct 22 (Figure 2) is provided with an entry end portion 36 having a thin wall tubular structure of circular cross-section (Figure 3). A second portion 38 of the duct 22 is welded to the ~t portion 36 and is pro-vided with a thin tubular wall which is formed with a circular cross-section at lts upper end. The duct por-tions 36 and 38 are generally inclined and shaped so that the duct channel extends somewhat radially inwardly along the turbine length.
A third or mouth duct portion 42 is provided with n outer tubular wall which is welded at its upstream end 43 to the exit end on the second duct portion 38. The tubular wall of the second and third duct portions 38 and ~0 42 is graduated in shape along its length from its up-stream circular cross~section to a cross-section at its downstream end 45 corresponding to a section of an annu-lus.
The mouth portion 42 is curved as indicated by the reference character 48 to direct the hot gases gener-ally in the longitudinal direction across the annular space through which the turbine blades rotate.
Heat from the hot transition duct wall metal is transferred to circulating coolant compressor air ~rom the outer duct wall surface. In addition, compressor coolant air is directed into a coolant channel network within the duct wall structure to provide efficient and enhanced wall metal cooling which leads to extended duct, blade and vane life.
It is preferred that at least the duct mouth portion 42 be provided with structure which provides internal wall cooling. If desired, other duct portions ~3~
6 48,931 can be provided with similar coolant provisions as war-ranted by tur~ine operating conditions.
Further, ducts provided with cooling features herein described can be provided in newly manufactured combustion turbines as well as retrofitted to turbines now operating in the ~ield.
As shown in Figures 7 and 8, the mouth duct por-tion 42 is provided with an outer shell 50 having a sup-port bracket 59A for attachment to the casing structure (Figure 1). An inwardly facing cooling cross channel 52 is provided about the duct mouth inner periphery at a point approximately midway along the length of the duct mouth portion 42. A plurality of coolant entry holes 57 are provided through the shell 50, to pass coolant air from the combustion casing inner space to the cross channel 52. The metal used for the duct shell 50, skin 54 and other duct wall structure can for example be either of the alloys previously noted herein.
An inner groove metallic skin 54 is attached, as by welding, to the inner surface of the shell 50. The outer surface of the skin 54 is provided with a plurality of outwardly facing coolant channels 56 which preferably extend parallel to each other in the longitudinal direc-tion to the upstream and downstream ends of the mouth duct portion. The channels 56 can for example be .03"
wide by .03" deep and spaced from each other by .03 inches. By "skin", it is meant to reer -to an element that is relatively thin so that good heat transfer is achieved for surface cooling.
All of the skin channels 56 are in communication with the transverse shell channel SZ to distribute coolant air, coming through the shell holes 57 from outside the duct, both upstream and downstream to the entry and dis-charge ends of the mouth portion 42. Accordingly, coolant air is substantially distri~uted about the entire duct wall periphery as it flows within the duct wall in the upstream and downstream directions.
As shown in Figure 4, coolant air is directly dis~harged from open downstream ends of the longitudinal ;, ~ ~3~
7 48,931 duct skin coolant channels 56 into the hot gas flo~, At the upstream end of the duct channels, an annular channel space 59 is provided between the end of the wall of the duct portlon 38 and the upstream ends of the longitudinal duct skin coolant channels 56 to provide for entry of the coolant into the downstream duct flow.
Calculated results achieved by use of the inven-tion in cooling a typical length of duct mouth are shown in Figure 9.
Since the inner tubular s~in surface is continu-ous and free of coolant outlet to the duct interior, it is readily adapted to accept a ceramic or other thermal bar-r~er coating. Further, undesirable temperature profiles across the duct flow are avoided to avoid life shortening effects on the turbine blades and vanes. Further, needed cool;ng action is achie~ed with comparatively less coolant flow thereby minimizing effects on turbine efficiency~
An alternate system for cooling the duct wall is to employ a transpirating material rather than the solid homogenous sXin 54 shown in Figures 6 and 8 with transpir-ating material 60 sho~n in Fi~res 10-15. Transpirating material such as "Lamilloy" ~ manufactured by the DDA
Div~sion of General Motors Corporation or similar laminates like Rolls Royce's "Transply" ~ may be used.
In this cooling arrangement, downward facing grooVes 61 are machined into the inner surface of shell 62, The sides 64 of the groove are canted to minimize the contact surface 65 with the laminated skin 60.
Air management is similar to that previously described except the channels serve as a distribution system delivering coolant air to the porous top surface of the ~aminate. The cooling air does not exit at the ends of coolant channels 63, but ~lows instead through the porous skin 600 The longitudinal grooves 61 terminate at the edge
Compressor discharge air provides external cooling ~or the tu~ular transition duct wall as it flows therearound as indicated by the illustrated ~low paths in Figure 1.
~ ~ ~3i~;i9~
48,931 ~ o provide high turbine operating efficiency and extended transition duct and blade and vane operating life, a transition duct is structured with enhanced wall cooling in accordance with the principles of the inven-tion. The preferred structure for the transition duct 22is shown in Figures 2-8.
The duct 22 (Figure 2) is provided with an entry end portion 36 having a thin wall tubular structure of circular cross-section (Figure 3). A second portion 38 of the duct 22 is welded to the ~t portion 36 and is pro-vided with a thin tubular wall which is formed with a circular cross-section at lts upper end. The duct por-tions 36 and 38 are generally inclined and shaped so that the duct channel extends somewhat radially inwardly along the turbine length.
A third or mouth duct portion 42 is provided with n outer tubular wall which is welded at its upstream end 43 to the exit end on the second duct portion 38. The tubular wall of the second and third duct portions 38 and ~0 42 is graduated in shape along its length from its up-stream circular cross~section to a cross-section at its downstream end 45 corresponding to a section of an annu-lus.
The mouth portion 42 is curved as indicated by the reference character 48 to direct the hot gases gener-ally in the longitudinal direction across the annular space through which the turbine blades rotate.
Heat from the hot transition duct wall metal is transferred to circulating coolant compressor air ~rom the outer duct wall surface. In addition, compressor coolant air is directed into a coolant channel network within the duct wall structure to provide efficient and enhanced wall metal cooling which leads to extended duct, blade and vane life.
It is preferred that at least the duct mouth portion 42 be provided with structure which provides internal wall cooling. If desired, other duct portions ~3~
6 48,931 can be provided with similar coolant provisions as war-ranted by tur~ine operating conditions.
Further, ducts provided with cooling features herein described can be provided in newly manufactured combustion turbines as well as retrofitted to turbines now operating in the ~ield.
As shown in Figures 7 and 8, the mouth duct por-tion 42 is provided with an outer shell 50 having a sup-port bracket 59A for attachment to the casing structure (Figure 1). An inwardly facing cooling cross channel 52 is provided about the duct mouth inner periphery at a point approximately midway along the length of the duct mouth portion 42. A plurality of coolant entry holes 57 are provided through the shell 50, to pass coolant air from the combustion casing inner space to the cross channel 52. The metal used for the duct shell 50, skin 54 and other duct wall structure can for example be either of the alloys previously noted herein.
An inner groove metallic skin 54 is attached, as by welding, to the inner surface of the shell 50. The outer surface of the skin 54 is provided with a plurality of outwardly facing coolant channels 56 which preferably extend parallel to each other in the longitudinal direc-tion to the upstream and downstream ends of the mouth duct portion. The channels 56 can for example be .03"
wide by .03" deep and spaced from each other by .03 inches. By "skin", it is meant to reer -to an element that is relatively thin so that good heat transfer is achieved for surface cooling.
All of the skin channels 56 are in communication with the transverse shell channel SZ to distribute coolant air, coming through the shell holes 57 from outside the duct, both upstream and downstream to the entry and dis-charge ends of the mouth portion 42. Accordingly, coolant air is substantially distri~uted about the entire duct wall periphery as it flows within the duct wall in the upstream and downstream directions.
As shown in Figure 4, coolant air is directly dis~harged from open downstream ends of the longitudinal ;, ~ ~3~
7 48,931 duct skin coolant channels 56 into the hot gas flo~, At the upstream end of the duct channels, an annular channel space 59 is provided between the end of the wall of the duct portlon 38 and the upstream ends of the longitudinal duct skin coolant channels 56 to provide for entry of the coolant into the downstream duct flow.
Calculated results achieved by use of the inven-tion in cooling a typical length of duct mouth are shown in Figure 9.
Since the inner tubular s~in surface is continu-ous and free of coolant outlet to the duct interior, it is readily adapted to accept a ceramic or other thermal bar-r~er coating. Further, undesirable temperature profiles across the duct flow are avoided to avoid life shortening effects on the turbine blades and vanes. Further, needed cool;ng action is achie~ed with comparatively less coolant flow thereby minimizing effects on turbine efficiency~
An alternate system for cooling the duct wall is to employ a transpirating material rather than the solid homogenous sXin 54 shown in Figures 6 and 8 with transpir-ating material 60 sho~n in Fi~res 10-15. Transpirating material such as "Lamilloy" ~ manufactured by the DDA
Div~sion of General Motors Corporation or similar laminates like Rolls Royce's "Transply" ~ may be used.
In this cooling arrangement, downward facing grooVes 61 are machined into the inner surface of shell 62, The sides 64 of the groove are canted to minimize the contact surface 65 with the laminated skin 60.
Air management is similar to that previously described except the channels serve as a distribution system delivering coolant air to the porous top surface of the ~aminate. The cooling air does not exit at the ends of coolant channels 63, but ~lows instead through the porous skin 600 The longitudinal grooves 61 terminate at the edge
3~ of t~e skin 60 (see Figures 13 and 14).
Claims (8)
1. A transition duct for a large plant combustion turbine comprising an elongated outer tubular wall member arranged to be coupled with a combustion chamber at its upstream end and to be supported at its downstream end to direct the hot driving gases through the annular turbine blade space, said tubular wall having a generally continuous inner surface free of film cooling and free of structural discontinuities otherwise created by slotting needed for such film cooling and having at least a portion thereof structured for internal wall cooling, said wall portion including an outer tubular shell member and a skin member having a continuous inner tubular surface forming a part of said continuous wall inner surface and an outer surface secured peripherally about and against an inward facing tubular surface of said shell member, coolant channel means arranged peripherally about and in said skin member to face said shell member and to direct coolant generally longitudinally of said duct portion between said shell and skin members, said shell member having entry means for directing compressor discharge coolant air from the duct exterior to and through said skin coolant channel means, and means for discharging the coolant from said channel means to the duct internal gas flow from at least one location longitudinally spaced from said entry means.
2. A transition duct as set forth in Claim 1 wherein said entry means comprises at least one substantially transverse coolant channel disposed peripherally about said tubular shell member to face inwardly toward said tubular skin member, and a plurality of openings extending from said shell coolant channel through said shell wall to the duct exterior and disposed periphery about said shell member.
3. A transition duct as set forth in claim 1 wherein said channel means comprises a plurality of sub-stantially parallel channels extending longitudinally between the upstream and downstream end of said duct portion, said discharge means comprising openings from said skin channels through the downstream end of said skin member.
4. A transition duct as set forth in claim 3 wherein said entry means comprises at least one substan-tially transverse coolant channel disposed peripherally about said tubular shell member to face inwardly toward said tubular skin member, and a plurality of openings extending from said shell coolant channel through said shell wall to the duct exterior and disposed periphery about said shell member;
said transverse shell channel is disposed sub-stantially midway between the upstream and downstream ends of said skin coolant channels, and said discharge means further includes openings from said skin channels through the upstream end of said skin member; and channel means facing the duct interior and extend-ing peripherally about said shell member at its upstream end to direct coolant air from the upstream ends of said skin channels into the duct gas flow.
said transverse shell channel is disposed sub-stantially midway between the upstream and downstream ends of said skin coolant channels, and said discharge means further includes openings from said skin channels through the upstream end of said skin member; and channel means facing the duct interior and extend-ing peripherally about said shell member at its upstream end to direct coolant air from the upstream ends of said skin channels into the duct gas flow.
5. A transition duct as set forth in claim 1 wherein the upstream end of said tubular duct has gener-ally circular cross-section and the downstream end of said tubular duct has a cross-section corresponding to a section of an annulus, and said internally cooled tubular wall port-ion is a downstream end portion of said duct where the annulus or near-annulus duct cross-section affects the duct strength against thermal stress and differential pressure.
6. A transition duct as set forth in claim 5 wherein said entry means comprises at least one substan-tially transverse coolant channel disposed peripherally about said tubular shell member to face inwardly toward said tubular skin member, and a plurality of openings extending from said shell coolant channel through said shell wall to the duct exterior and diposed periphery about said shell member; and said channel means comprises a plurality of sub-stantially parallel channels extending longitudinally be-tween the upstream and downstream ends of said internally cooled turbine wall portion, said discharge means comprising openings from said skin channels through the downstream end of said skin member.
7. An assembly for a large plant combustion turbine comprising at least one generally cylindrical combustion chamber, a plurality of turbine blades disposed to be rotated through an annular space under the driving force of hot gas, a transition duct having a tubular wall with an upstream section having a generally circular cross-section and coupled to said combustion chamber and with a downstream section have a discharge end cross-section corresponding to a section of an annulus to direct the hot driving gases through the turbine blade annulus, means for supporting said transition duct, said tubular wall having a generally continuous inner surface free of film cooling and free of structural discontinuities otherwise created by slotting needed for such film cooling and having at least a portion thereof structured for internal wall cooling, said wall portion including an outer tubular shell member and a skin member having a continuous inner tubular surface forming a part of said continuous wall inner surface and an outer surface secured peripherally about and against an in-ward facing tubular surface of said shell member, coolant channel means arranged peripherally about and in said skin member to face said shell member and to direct coolant generally longitudinally of said duct portion between said shell and skin members, said shell member having entry means for directing compressor discharge coolant air from the duct exterior to and through said skin coolant channel means, and means for discharging the coolant from said channel means to the duct internal gas flow from at least one location longitudinally spaced from said entry means.
8. A turbine assembly as set forth in claim 7 wherein said duct portion is said downstream end section where the annulus or near-annulus duct cross-section affects the duct strength against thermal stress and differential pressure.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US20444180A | 1980-11-06 | 1980-11-06 | |
| US204,441 | 1980-11-06 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1183695A true CA1183695A (en) | 1985-03-12 |
Family
ID=22757890
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000388209A Expired CA1183695A (en) | 1980-11-06 | 1981-10-19 | Efficiently cooled transition duct for a large plant combustion turbine |
Country Status (8)
| Country | Link |
|---|---|
| JP (1) | JPS6027816B2 (en) |
| AR (1) | AR225977A1 (en) |
| BE (1) | BE891023A (en) |
| BR (1) | BR8106793A (en) |
| CA (1) | CA1183695A (en) |
| GB (1) | GB2087066B (en) |
| IT (1) | IT1142046B (en) |
| MX (1) | MX154157A (en) |
Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH0752014B2 (en) * | 1986-03-20 | 1995-06-05 | 株式会社日立製作所 | Gas turbine combustor |
| US6018950A (en) * | 1997-06-13 | 2000-02-01 | Siemens Westinghouse Power Corporation | Combustion turbine modular cooling panel |
| CA2288557C (en) * | 1998-11-12 | 2007-02-06 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor cooling structure |
| US6412268B1 (en) * | 2000-04-06 | 2002-07-02 | General Electric Company | Cooling air recycling for gas turbine transition duct end frame and related method |
| US7310938B2 (en) * | 2004-12-16 | 2007-12-25 | Siemens Power Generation, Inc. | Cooled gas turbine transition duct |
| US8015818B2 (en) | 2005-02-22 | 2011-09-13 | Siemens Energy, Inc. | Cooled transition duct for a gas turbine engine |
| JP2010085052A (en) * | 2008-10-01 | 2010-04-15 | Mitsubishi Heavy Ind Ltd | Combustor tail pipe, designing method therefor, and gas turbine |
-
1981
- 1981-10-08 GB GB8130376A patent/GB2087066B/en not_active Expired
- 1981-10-13 MX MX189619A patent/MX154157A/en unknown
- 1981-10-19 CA CA000388209A patent/CA1183695A/en not_active Expired
- 1981-10-22 BR BR8106793A patent/BR8106793A/en unknown
- 1981-10-30 AR AR287296A patent/AR225977A1/en active
- 1981-11-05 IT IT24871/81A patent/IT1142046B/en active
- 1981-11-06 JP JP56177265A patent/JPS6027816B2/en not_active Expired
- 1981-11-06 BE BE0/206466A patent/BE891023A/en not_active IP Right Cessation
Also Published As
| Publication number | Publication date |
|---|---|
| IT1142046B (en) | 1986-10-08 |
| IT8124871A0 (en) | 1981-11-05 |
| BR8106793A (en) | 1982-07-06 |
| BE891023A (en) | 1982-05-06 |
| MX154157A (en) | 1987-05-27 |
| JPS6027816B2 (en) | 1985-07-01 |
| JPS57113923A (en) | 1982-07-15 |
| GB2087066B (en) | 1984-09-19 |
| GB2087066A (en) | 1982-05-19 |
| AR225977A1 (en) | 1982-05-14 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| MKEX | Expiry |