CA1183694A - Efficiently cooled combustor for a combustion turbine - Google Patents
Efficiently cooled combustor for a combustion turbineInfo
- Publication number
- CA1183694A CA1183694A CA000402712A CA402712A CA1183694A CA 1183694 A CA1183694 A CA 1183694A CA 000402712 A CA000402712 A CA 000402712A CA 402712 A CA402712 A CA 402712A CA 1183694 A CA1183694 A CA 1183694A
- Authority
- CA
- Canada
- Prior art keywords
- combustor
- shell
- skin
- coolant
- ring
- 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
- 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)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
ABSTRACT OF THE DISCLOSURE
A combustor for a combustion turbine comprises a plurality of ring segments, each with an enhanced wall cooling mechanism. Each ring segment comprises a shell member and a skin member with coolant channels spiraling longitudinally therebetween for conduction of cooling air from the ring segment exterior at one end to the segment interior at the opposite end.
A combustor for a combustion turbine comprises a plurality of ring segments, each with an enhanced wall cooling mechanism. Each ring segment comprises a shell member and a skin member with coolant channels spiraling longitudinally therebetween for conduction of cooling air from the ring segment exterior at one end to the segment interior at the opposite end.
Description
3~
1 49,3g4 EFFICIENTLY COOLED COMBUSTOR
FOR A COMBUSTION TURBINE
BACKGROUND OF THE INVENTION
The present invention relates to combustion turbines as may be employed in a variety of uses, such as industrial processes, electric power generation, or air-craft engines. More particularly, the present inventionis directed to combustors employed in combustion turbines to heat motive gases which drive the turbine. The des-cription hereafter details the structure of a combustion turbine used for power generation, but applies as well ~o any type of combustion turbine.
A combustion turbine used for electric power generation typically has some eight to sixteen combustors peripherally arranged about the turbine longitudinal axis.
Combustion of fuel within each combustor heats pressurized 15 air supplied from a compressor section of the turbine.
The hot motive gas generated within ~ach combustor flows by means of a transition duct to a turbine section of the combustion turbine. Impingement of the hot motive gases on turbine blades causes rotation of 'che turbine rotor and generation of power.
Continuing ~fforts to obtain higher power gener-ation at greater efficiency have resulted in a continuous increase in gas operating temperatures. Even with the use of impro~d stainless steel or nickel-based alloys such as Hastalloy~or Inco ~ 7, it has become increasingly diffi-cult to provide ade~uate cooling of combustors so as to "~ ' , . .
1 49,3g4 EFFICIENTLY COOLED COMBUSTOR
FOR A COMBUSTION TURBINE
BACKGROUND OF THE INVENTION
The present invention relates to combustion turbines as may be employed in a variety of uses, such as industrial processes, electric power generation, or air-craft engines. More particularly, the present inventionis directed to combustors employed in combustion turbines to heat motive gases which drive the turbine. The des-cription hereafter details the structure of a combustion turbine used for power generation, but applies as well ~o any type of combustion turbine.
A combustion turbine used for electric power generation typically has some eight to sixteen combustors peripherally arranged about the turbine longitudinal axis.
Combustion of fuel within each combustor heats pressurized 15 air supplied from a compressor section of the turbine.
The hot motive gas generated within ~ach combustor flows by means of a transition duct to a turbine section of the combustion turbine. Impingement of the hot motive gases on turbine blades causes rotation of 'che turbine rotor and generation of power.
Continuing ~fforts to obtain higher power gener-ation at greater efficiency have resulted in a continuous increase in gas operating temperatures. Even with the use of impro~d stainless steel or nickel-based alloys such as Hastalloy~or Inco ~ 7, it has become increasingly diffi-cult to provide ade~uate cooling of combustors so as to "~ ' , . .
2 49,394 assure long combustor life. At present, the temperature of operating gases is typically 2300F while the tempera-ture operating limit for special metals used in construc-tion of combustors is about 1500F. It is expected that turbine gas operating temperatures will increase even further in the future.
Typical prior art combustors are presently cooled by introduction of air in tha form of a thin film along an interior surface of a ~ombustor wall. The pri-mary purpose of the film of air is to prevent impingementof hot motive gas upon the interior surace of the com-bustor wall. Such interior film cooling has generally been adequate. However, it is becoming less so with increasing turbine qas operating temperatures for three reasons. First, depletion of the cooling film with hot motive gas tends to defeat the insulating properties of the cooling film. Second, inability of the cooling air to persist as a jet film due to the force of impingement by the hot mot:i~e gases destroys the film effect and reduces the effectiveness of th~ cooling techniqu~. Finally, the use of fuels which generate a highly radiative flame increases the heat load on the combustor wall.
A second prior art combustor cooling technique involves use of a porous, laminated liner within the combustor, both to cool the liner as well as to provide a film of cooling air between the liner and the hot motive gases. The laminated liner comprises a plurality layers, each having channels formed therein, the channels of adjacent layers oriented generally perpendicular to one another. The channels of adjacent layers oriented gener-ally perpendicular to one another. The channels of adj-acent layers are interconnected by means of holes in the layers to provide flow communication through the liner to the interior of the combustor.
Transpiration cooling has also generally been adequate. However, increasing gas operating temperatures has resulted in the same type of problems for a trans-
Typical prior art combustors are presently cooled by introduction of air in tha form of a thin film along an interior surface of a ~ombustor wall. The pri-mary purpose of the film of air is to prevent impingementof hot motive gas upon the interior surace of the com-bustor wall. Such interior film cooling has generally been adequate. However, it is becoming less so with increasing turbine qas operating temperatures for three reasons. First, depletion of the cooling film with hot motive gas tends to defeat the insulating properties of the cooling film. Second, inability of the cooling air to persist as a jet film due to the force of impingement by the hot mot:i~e gases destroys the film effect and reduces the effectiveness of th~ cooling techniqu~. Finally, the use of fuels which generate a highly radiative flame increases the heat load on the combustor wall.
A second prior art combustor cooling technique involves use of a porous, laminated liner within the combustor, both to cool the liner as well as to provide a film of cooling air between the liner and the hot motive gases. The laminated liner comprises a plurality layers, each having channels formed therein, the channels of adjacent layers oriented generally perpendicular to one another. The channels of adjacent layers oriented gener-ally perpendicular to one another. The channels of adj-acent layers are interconnected by means of holes in the layers to provide flow communication through the liner to the interior of the combustor.
Transpiration cooling has also generally been adequate. However, increasing gas operating temperatures has resulted in the same type of problems for a trans-
3 49,394 piration-cooled combustor as were described above for a film-cooled combustor. In addition, transpiration cooling requires significant volumes of compr~ssor discharge air to provide ade~uate cooling, depleting the air which would normally be used to drive thP turbine end thereby reducing turbine efficiency.
Hence, it would be advantageous to develop a cooling apparatus which provides improved cooling effi-ciency for a combustor wall, thereby permitting hisner gas operating temperatures and resulting in longer combustor life, and reducing the volume of cooling air drawn from the compressor section, permitting higher turbine operat-ing efficiency.
SU~ ~ RY OF THE INVENTION
Accordingly, a combustor for a combustion tur-bine comprises a generally cylindrical combustor preer-ably formed from a plurality of ring segments coupled together with a conical segment at a fuel-injection end of the combustor. Each segment comprises an outer shell member and an inner skin member, a combination thereof deining coolant channel means for dirscting flow of cooling air. Cooling air enters from the combustor exter-ior at one end of each segment through the shell member, flows through the channel means between the shell and skin members removing heat therefrom, and exits to the interior of the combustor at the opposite end of the ring segment.
BhlEE DESCRIPI IN OF THE DRAWINGS
Fig. 1 shows a longitudinal section of a combus-tion turbine in which a combustor is arranged to be cooled in accordance with the principles of the invention;
Fig. 2 shows an elevational view of an upper portion of the comh~lstor along a longitudinal section thereof;
Fig. 3 shows a ring subassembly of the combustor in section;
Fig. 4 shows an upstream end view o a cross-~ection of the ring subasse~bly at a coolant inlet hole;
~ ~3~
Hence, it would be advantageous to develop a cooling apparatus which provides improved cooling effi-ciency for a combustor wall, thereby permitting hisner gas operating temperatures and resulting in longer combustor life, and reducing the volume of cooling air drawn from the compressor section, permitting higher turbine operat-ing efficiency.
SU~ ~ RY OF THE INVENTION
Accordingly, a combustor for a combustion tur-bine comprises a generally cylindrical combustor preer-ably formed from a plurality of ring segments coupled together with a conical segment at a fuel-injection end of the combustor. Each segment comprises an outer shell member and an inner skin member, a combination thereof deining coolant channel means for dirscting flow of cooling air. Cooling air enters from the combustor exter-ior at one end of each segment through the shell member, flows through the channel means between the shell and skin members removing heat therefrom, and exits to the interior of the combustor at the opposite end of the ring segment.
BhlEE DESCRIPI IN OF THE DRAWINGS
Fig. 1 shows a longitudinal section of a combus-tion turbine in which a combustor is arranged to be cooled in accordance with the principles of the invention;
Fig. 2 shows an elevational view of an upper portion of the comh~lstor along a longitudinal section thereof;
Fig. 3 shows a ring subassembly of the combustor in section;
Fig. 4 shows an upstream end view o a cross-~ection of the ring subasse~bly at a coolant inlet hole;
~ ~3~
4 ~g,394 Fig. 5 shows an upstream end view of a cross-section o the ring subassembly at a point downstream of the coolant inlet hole;
Fig. 6 shows a top view of a cross section of the ring subassembly depicting spiral channels within a skin member;
Fig. 7 shows an enlarged view of a portion of Fig. 5 with an alternative configuration of channelsi Fig. 8 shows an alternative embodiment of the ring subassembly in section.
DESCRIPTION OF THE PREFERRED EMBODIMENT
More particularly, there is shown in Fig. 1 a combustion turbine 10 having a plurality of generally cylindrical combustors 12. Fuel is supplied to the com-bustors 12 through a nozzle structure 14 and air is sup-plied to the combustors 12 by a compressor 16 through air flow space 18 within a combustion casing 23.
Hot gaseous products of combustion are directed from each combustor 12 through a transition duct 22 where they are discharged into the annular space through which turbine blades 24, 26 rotate under the driving force of the expanding gases.
To provide high turbine operating efficiency and extended combustor operating life, a combustor is ætruc-tured with enhanced wall cooling in accordance with theprinciples of the invention. The preferred structure for the combustor 12 is shown in Figs. 2 8.
The combustor 12 (Fig. 2) compri~es a plurality of cylindrical ring subassemblies 30 and a conical ring subassembly 32 all joined by appropriate means such as welding. The lengkh of a single ring subass~mbly 30, 32 is shown as extending between reference numbers 34 and 36.
Cylindrical ring subassemblies 30 are joined end-to-end to form a cylindrical combustor assembly 12. The conical ring subassembly 32 is attached to the combustor assembly 12 at a fuel intake end of the assembly 12. Each ring subassembly 30, 32 comprises a cooling mechanism complete ~3~
4~,394 within itself but cooperating with the cooling mechanism of every other ring subassembly 30, 32 to provide an efficiently cooled combustor assembly 12.
Fig. 3 shows a ring subassembly 30 of the com-bustor 12 in section. Structure of a conical æubassembly32 is substantially like that described below for a cyl-indrical subassembly 30. Each ring subassembly 30 com-prises a cylindrical exterior shell 38 surrounding and adjoined to a cylindrical interior skin 40. The exterior shell 38 is continuous from one end of the ring sub~
assembly 30 to the other end about the entire exterior surface of the subassembly 30 except for a plurality of coolant inlet holes 42 positioned around the circumference of the subassembly 30 at the end of each subassembly 30 nearest the fuel intake end of the combustor assembly 12.
The inlet holes 42 provide an entrance for cooling air 44 into a cross channel 46 defined by an inward facing groove in the shell 40 and extending around the circumferenc~ of the ring subassembly 30. Fig. 4 de-picts the inlet holes 42 and the cross channel 46 from adifferent view. The cross channel 46 is in flow communi-cation with a plurality of coolant channels 48 between the shell 38 and the skin 40. The coolant channels 48 shown in greater detail in Figs. 4 through 7, extend from one end of the subassembly 30 to the opposite end.
Fig. 6 depicts spiral grooves 50 which are formed in an outer surface of the skin 40 by an appropri-ate technique such as machining or etching. The grooved surface of the skin 40 is attached to khe inner surface of the shell 38 to define parallel coolant channels 48 which æpiral longitudinally about the circumference of the æubassembly 30 from the cross channel 46 to khe exit point 52. The coolant channels 48 can, for example, be .03 inches wide by .03 inches deep and spaced from each other by .03 inches.
An alternative to the rectangular cross æec-tional structure of coolant channels 48 is depicted in 3~g~
6 49,3~4 Fig. 7. In this arrangement, the coolant channels 48 are formed by grooving the interior surface of the shell 38.
The walls 54 of the channels 48 are canted to provide a decrease in actual surface contact between the shell 38 and the skin 40 and a resultant increase in surface con~
tact between the skin 38 and cooling air within the cool-ant channels 48. Canting the channel walls 54 also in-creases surface contact between the channel walls 54 and the cooling air. The canted channel arrangement thus pro~
vides for more efficient transfer of h~at from the inter-ior skin 40 and exterior shell 38 to the cooling air within the channel 48.
The spiral arrangement of coolant channels 48 is preferred over a strictly longitudinal arrangement.
Spiraling the channels 4~ increases the effective length of the channels 48 without increasing the length of the ring subassembly 30. The length of a spiral channel 48 may be controlled by the angle chosen for the spiral.
Spiraling the channels 48 also improves the distribution of coolant channels 48 within the subassembly 30. For example, the spiral channels 48 more effectively protect the combustor assembly 12 from a streak of hot motive gas flowing in a straight line through the combustor assembly 12 by providing the cooling effect of a plurality of cooling channels 48 as opposed to the cooling effect of a few longitudinal channels.
Control of cooling air flow within a ring sub-asse~bly 30 may be accomplished by variation of fiv~
structural features: (l) the cross-sectional dimensions of a coolant channel 48; (2) the number of coolant channels 48; (3) the angle of spiral of the channels 48; (4) the length of the ring subassembly 30; and ~5) the dimensions of the inlet holes. Each ring subassembly 30 may be individually designed to achieve the cooling characteris-35 tics required for its position within the combustor assem-bly .
3~
7 ~9,394 An alternative structure o the ring subassambly 30 is depicted in Fig. 8. Each ring subassembly 30 com-prises an exterior shell 38, an interior skin 40 and an interior thermal barrier 60 bonded to the inner surface of the skin 40. The relatively short leng~h of the ring subassembly 30 and the non-porous nature of the skin 40 permits easy application of the thermal barrier 60 to the interior of the combustor 12. The thermal barrier 60, typically 0.015 to 0.025 inches in depth, may be any ceramic or other thermal barrier coating, such as yttrium-stabilized zirconium oxide, which effectively insulates the skin 40 and shell 38 of the ring subassambly 30 from the harsh interior temperatures of the combustor 12. The thermal barr.ier 60 provides a feature with struc-tural characteristics such as material type, grade andthickness which may be varied with each subassembly to achieve desired insulating characteristics.
Fig. 6 shows a top view of a cross section of the ring subassembly depicting spiral channels within a skin member;
Fig. 7 shows an enlarged view of a portion of Fig. 5 with an alternative configuration of channelsi Fig. 8 shows an alternative embodiment of the ring subassembly in section.
DESCRIPTION OF THE PREFERRED EMBODIMENT
More particularly, there is shown in Fig. 1 a combustion turbine 10 having a plurality of generally cylindrical combustors 12. Fuel is supplied to the com-bustors 12 through a nozzle structure 14 and air is sup-plied to the combustors 12 by a compressor 16 through air flow space 18 within a combustion casing 23.
Hot gaseous products of combustion are directed from each combustor 12 through a transition duct 22 where they are discharged into the annular space through which turbine blades 24, 26 rotate under the driving force of the expanding gases.
To provide high turbine operating efficiency and extended combustor operating life, a combustor is ætruc-tured with enhanced wall cooling in accordance with theprinciples of the invention. The preferred structure for the combustor 12 is shown in Figs. 2 8.
The combustor 12 (Fig. 2) compri~es a plurality of cylindrical ring subassemblies 30 and a conical ring subassembly 32 all joined by appropriate means such as welding. The lengkh of a single ring subass~mbly 30, 32 is shown as extending between reference numbers 34 and 36.
Cylindrical ring subassemblies 30 are joined end-to-end to form a cylindrical combustor assembly 12. The conical ring subassembly 32 is attached to the combustor assembly 12 at a fuel intake end of the assembly 12. Each ring subassembly 30, 32 comprises a cooling mechanism complete ~3~
4~,394 within itself but cooperating with the cooling mechanism of every other ring subassembly 30, 32 to provide an efficiently cooled combustor assembly 12.
Fig. 3 shows a ring subassembly 30 of the com-bustor 12 in section. Structure of a conical æubassembly32 is substantially like that described below for a cyl-indrical subassembly 30. Each ring subassembly 30 com-prises a cylindrical exterior shell 38 surrounding and adjoined to a cylindrical interior skin 40. The exterior shell 38 is continuous from one end of the ring sub~
assembly 30 to the other end about the entire exterior surface of the subassembly 30 except for a plurality of coolant inlet holes 42 positioned around the circumference of the subassembly 30 at the end of each subassembly 30 nearest the fuel intake end of the combustor assembly 12.
The inlet holes 42 provide an entrance for cooling air 44 into a cross channel 46 defined by an inward facing groove in the shell 40 and extending around the circumferenc~ of the ring subassembly 30. Fig. 4 de-picts the inlet holes 42 and the cross channel 46 from adifferent view. The cross channel 46 is in flow communi-cation with a plurality of coolant channels 48 between the shell 38 and the skin 40. The coolant channels 48 shown in greater detail in Figs. 4 through 7, extend from one end of the subassembly 30 to the opposite end.
Fig. 6 depicts spiral grooves 50 which are formed in an outer surface of the skin 40 by an appropri-ate technique such as machining or etching. The grooved surface of the skin 40 is attached to khe inner surface of the shell 38 to define parallel coolant channels 48 which æpiral longitudinally about the circumference of the æubassembly 30 from the cross channel 46 to khe exit point 52. The coolant channels 48 can, for example, be .03 inches wide by .03 inches deep and spaced from each other by .03 inches.
An alternative to the rectangular cross æec-tional structure of coolant channels 48 is depicted in 3~g~
6 49,3~4 Fig. 7. In this arrangement, the coolant channels 48 are formed by grooving the interior surface of the shell 38.
The walls 54 of the channels 48 are canted to provide a decrease in actual surface contact between the shell 38 and the skin 40 and a resultant increase in surface con~
tact between the skin 38 and cooling air within the cool-ant channels 48. Canting the channel walls 54 also in-creases surface contact between the channel walls 54 and the cooling air. The canted channel arrangement thus pro~
vides for more efficient transfer of h~at from the inter-ior skin 40 and exterior shell 38 to the cooling air within the channel 48.
The spiral arrangement of coolant channels 48 is preferred over a strictly longitudinal arrangement.
Spiraling the channels 4~ increases the effective length of the channels 48 without increasing the length of the ring subassembly 30. The length of a spiral channel 48 may be controlled by the angle chosen for the spiral.
Spiraling the channels 48 also improves the distribution of coolant channels 48 within the subassembly 30. For example, the spiral channels 48 more effectively protect the combustor assembly 12 from a streak of hot motive gas flowing in a straight line through the combustor assembly 12 by providing the cooling effect of a plurality of cooling channels 48 as opposed to the cooling effect of a few longitudinal channels.
Control of cooling air flow within a ring sub-asse~bly 30 may be accomplished by variation of fiv~
structural features: (l) the cross-sectional dimensions of a coolant channel 48; (2) the number of coolant channels 48; (3) the angle of spiral of the channels 48; (4) the length of the ring subassembly 30; and ~5) the dimensions of the inlet holes. Each ring subassembly 30 may be individually designed to achieve the cooling characteris-35 tics required for its position within the combustor assem-bly .
3~
7 ~9,394 An alternative structure o the ring subassambly 30 is depicted in Fig. 8. Each ring subassembly 30 com-prises an exterior shell 38, an interior skin 40 and an interior thermal barrier 60 bonded to the inner surface of the skin 40. The relatively short leng~h of the ring subassembly 30 and the non-porous nature of the skin 40 permits easy application of the thermal barrier 60 to the interior of the combustor 12. The thermal barrier 60, typically 0.015 to 0.025 inches in depth, may be any ceramic or other thermal barrier coating, such as yttrium-stabilized zirconium oxide, which effectively insulates the skin 40 and shell 38 of the ring subassambly 30 from the harsh interior temperatures of the combustor 12. The thermal barr.ier 60 provides a feature with struc-tural characteristics such as material type, grade andthickness which may be varied with each subassembly to achieve desired insulating characteristics.
Claims (7)
1. A combustor for a combustion turbine comprising a plurality of ring members coupled to form an elongated chamber wherein compressed gases are heated for driving a turbine, each of said ring members comprising an outer shell member and an inner skin member, means forming a part of said shell and skin members defining coolant channel means arranged peripherally about and between said shell and skin members to direct coolant generally longitudinally of said ring member between said shell and skin members, said coolant channel means comprising a plurality of generally parallel channels spiraling about the longitudinal axis of said ring member within said combination of said shell and skin members, said shell member having entry means for direct-ing compressor discharge coolant air from the combustor exterior to and through said spiraling channels, and said shell and skin members having means for discharging the coolant from said spiraling channel to a combustor internal gas flow.
2. A combustor as set forth in claim 1 wherein said entry means comprises a plurality of peripheral coolant inlet holes adjacent an end of said ring member in flow communication through an annular cross channel means with said coolant channel means, said cross channel means comprising a circumferential inwardly facing groove in said shell member subadjacent the inlet holes.
3. A combustor as set forth in claim 1 wherein said spiraling channels are defined by outwardly facing grooves on the outwardly facing surface of said skin member.
4. A combustor as set forth in claim 1 wherein said discharge means comprises an exit channel providing flow communication between said coolant channel means and the combustor interior, the exit channel defined by an extension of said shell member beyond the end of said skin member leaving said coolant channel means open to the combustor interior.
5. A combustor as set forth in claim 1 wherein the combustor includes a conical ring member at a fuel-injection end of the combustor.
6. A combustor as set forth in claim 1 wherein said coolant channel means comprises a plurality of gener-ally parallel channels spiraling about the longitudinal axis of said ring member within said combination of said shell and skin members, the channels defined by inwardly facing grooves in the surface of said shell member, each such groove having walls canted to define a groove base narrower than a groove mouth.
7. A combustor as set forth in claim 1 includ-ing a thermal barrier member bonded to and continuous with the inner side of the skin member.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US27285181A | 1981-06-12 | 1981-06-12 | |
| US272,851 | 1981-06-12 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1183694A true CA1183694A (en) | 1985-03-12 |
Family
ID=23041580
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000402712A Expired CA1183694A (en) | 1981-06-12 | 1982-05-11 | Efficiently cooled combustor for a combustion turbine |
Country Status (8)
| Country | Link |
|---|---|
| JP (1) | JPS582528A (en) |
| AR (1) | AR227595A1 (en) |
| BE (1) | BE893491A (en) |
| BR (1) | BR8203361A (en) |
| CA (1) | CA1183694A (en) |
| GB (1) | GB2102558B (en) |
| IT (1) | IT1151263B (en) |
| MX (1) | MX158473A (en) |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE3615226A1 (en) * | 1986-05-06 | 1987-11-12 | Mtu Muenchen Gmbh | HOT GAS OVERHEATING PROTECTION DEVICE FOR GAS TURBINE ENGINES |
| DE59010740D1 (en) * | 1990-12-05 | 1997-09-04 | Asea Brown Boveri | Gas turbine combustor |
| US5239832A (en) * | 1991-12-26 | 1993-08-31 | General Electric Company | Birdstrike resistant swirler support for combustion chamber dome |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4253301A (en) * | 1978-10-13 | 1981-03-03 | General Electric Company | Fuel injection staged sectoral combustor for burning low-BTU fuel gas |
-
1982
- 1982-05-11 CA CA000402712A patent/CA1183694A/en not_active Expired
- 1982-05-20 MX MX192793A patent/MX158473A/en unknown
- 1982-06-02 IT IT21634/82A patent/IT1151263B/en active
- 1982-06-04 GB GB08216384A patent/GB2102558B/en not_active Expired
- 1982-06-08 BR BR8203361A patent/BR8203361A/en unknown
- 1982-06-11 BE BE0/208330A patent/BE893491A/en not_active IP Right Cessation
- 1982-06-11 JP JP57099413A patent/JPS582528A/en active Pending
- 1982-07-10 AR AR289660A patent/AR227595A1/en active
Also Published As
| Publication number | Publication date |
|---|---|
| IT1151263B (en) | 1986-12-17 |
| IT8221634A0 (en) | 1982-06-02 |
| AR227595A1 (en) | 1982-11-15 |
| BR8203361A (en) | 1984-01-10 |
| MX158473A (en) | 1989-02-03 |
| JPS582528A (en) | 1983-01-08 |
| GB2102558B (en) | 1985-02-13 |
| BE893491A (en) | 1982-12-13 |
| GB2102558A (en) | 1983-02-02 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| MKEC | Expiry (correction) | ||
| MKEX | Expiry |