EP3921577A1 - Rohrbrennkammersystem und gasturbinenanlage mit einem solchen rohrbrennkammersystem - Google Patents
Rohrbrennkammersystem und gasturbinenanlage mit einem solchen rohrbrennkammersystemInfo
- Publication number
- EP3921577A1 EP3921577A1 EP20711063.6A EP20711063A EP3921577A1 EP 3921577 A1 EP3921577 A1 EP 3921577A1 EP 20711063 A EP20711063 A EP 20711063A EP 3921577 A1 EP3921577 A1 EP 3921577A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- combustion chamber
- turbine
- chamber system
- tubular combustion
- lining
- 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.)
- Granted
Links
- 238000002485 combustion reaction Methods 0.000 title claims abstract description 38
- 230000007704 transition Effects 0.000 claims abstract description 51
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 6
- 239000000919 ceramic Substances 0.000 claims description 6
- 239000002184 metal Substances 0.000 claims description 6
- 229910052751 metal Inorganic materials 0.000 claims description 6
- 230000002093 peripheral effect Effects 0.000 claims description 4
- 239000007769 metal material Substances 0.000 claims description 3
- 238000001816 cooling Methods 0.000 description 7
- 239000000463 material Substances 0.000 description 4
- 238000012958 reprocessing Methods 0.000 description 3
- 238000007789 sealing Methods 0.000 description 3
- 230000008646 thermal stress Effects 0.000 description 3
- PXHVJJICTQNCMI-UHFFFAOYSA-N Nickel Chemical compound [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 description 2
- 238000013016 damping Methods 0.000 description 2
- 238000012423 maintenance Methods 0.000 description 2
- 230000008092 positive effect Effects 0.000 description 2
- 239000007787 solid Substances 0.000 description 2
- 238000005299 abrasion Methods 0.000 description 1
- 239000000654 additive Substances 0.000 description 1
- 230000000996 additive effect Effects 0.000 description 1
- 230000032683 aging Effects 0.000 description 1
- 229910010293 ceramic material Inorganic materials 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000003779 heat-resistant material Substances 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 238000009413 insulation Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 229910001092 metal group alloy Inorganic materials 0.000 description 1
- 229910052759 nickel Inorganic materials 0.000 description 1
- 238000003825 pressing Methods 0.000 description 1
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
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
-
- 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
- F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
-
- 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
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/46—Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings
Definitions
- Tube combustion chamber system and gas turbine system with one
- the present invention relates to a tubular combustion chamber system for a gas turbine system with a plurality of ring-shaped arranged th transition lines, which are designed to be connected with their upstream ends to a burner and to direct hot gas generated by the burner to a turbine. Furthermore, the present invention relates to a gas turbine system with several annularly arranged burners, a turbine and a tubular combustion chamber system of the type described above, which connects the burner to the turbine.
- Tube combustion chamber systems of the type mentioned at the beginning are used in gas turbine systems to convey hot gas from the burners to the turbine inlet.
- they comprise transition lines designed as pipelines, which are also referred to as "transitions" in specialist circles.
- the transition lines are thermally highly stressed during operation of the gas turbine system. Accordingly, they are made of high-temperature-resistant materials internal cooling channels and an internal layer system for thermal insulation (TBC + MCrAlY).
- Sealing systems are provided in the area of the interface to the turbine inlet to reduce the leakage of compressed air into the combustion system and to reduce relative movements between the tubular combustion chamber system and the turbine as well as between the Due to the design of the sealing systems and the mechanical degree of freedom of the interface between the transition lines and the turbine, on the one hand the side seals are subject to severe abrasion wear. On the other hand, the transition lines and their inner layer system also wear out due to the high thermal load, primarily in the exit area as a result of layer aging and sealing groove wear. Come in addition, that the flow to the turbine is not uniform due to the circumferentially discontinuous inflow cross-section at the interface between the transition lines and the turbine. The uneven flow as a result of the shading through the side walls and seals of the exit area of the transition lines causes high-frequency temperature and speed changes with negative effects on the service life of the turbine blades.
- the service life of the transition lines is determined by the
- the reconditioning also includes the stripping of the entire layer system and the recoating. The costs of this complex reprocessing are therefore close to the costs of new parts.
- the life cycle costs of new or existing gas turbine systems are primarily determined by the service life and maintenance intervals of the hot gas components.
- considerably longer maintenance intervals are required for new gas turbine systems while at the same time increasing thermal stress.
- constructive solutions are required that eliminate or at least significantly improve the weak points of current designs.
- an object of the present invention is to create a tubular combustion chamber system of the type mentioned at the outset with an improved design.
- the present invention creates a tubular combustion chamber system of the type mentioned, which is characterized in that it has a hot gas distributor designed for connection to the turbine, which defines an annular channel open to the turbine, into which the downstream ends of the transition lines open .
- a hot gas distributor designed for connection to the turbine, which defines an annular channel open to the turbine, into which the downstream ends of the transition lines open .
- Such an additional hot gas distributor between the transition lines and the turbine inlet leads to a very uniform flow to the turbine, which significantly reduces high-frequency temperature and speed changes. This has a very positive effect on the service life of the turbine blades.
- the transition lines and the hot gas distributor are made of metal and are provided on the inside with a refractory lining, in particular with a ceramic lining. Thanks to such a lining, the thermal stress on the metallic components, i.e. the hot gas distributor and the transition lines, is significantly reduced. The associated lower expansion differences in the area of the seals to the turbine and the seals between the transition lines lead to less wear in this area and enable more solid joining concepts between the tubular combustion chamber system and the turbine. In addition, the refractory lining entails lower high temperature requirements for the materials of the metallic components, which can save costs.
- the transition lines can be designed without an internal layer system, which significantly reduces the effort required for repair and reconditioning, as there is no need to decoat and recoat the transition lines.
- the use of a refractory lining also reduces the cooling requirement for the metallic components of the tubular combustion chamber system. Compared to tubular combustion chamber systems without ceramic lining, the cooling air requirement can, according to current calculations, be up to 50%, which leads to an increase in the performance of the gas turbine system.
- each transition line tapers conically in the downstream direction, the refractory lining of the transition line having at least one annular lining section with a conically tapering outside diameter in the downstream direction, which is held on the transition line with radial and axial bias.
- a bias which can be realized for example by positioning spring and / or damping elements between the refractory lining and the corresponding transition line, differences in thermal expansion between the metallic transition lines and their ceramic lining are compensated.
- the ceramic lining is fixed to a limited extent under all operating conditions.
- the at least one annular lining section can be formed by a single lining element, that is to say by an annular lining element with a conical outer surface.
- the at least one ring-shaped lining section by several ring-segment-shaped lining elements which are braced against one another in the circumferential direction.
- the refractory lining of the hot gas distributor has in front of geous a variety of lining elements that are attached to the radially inside and outside surfaces of the hot gas distributor un ter radial bias.
- the installation of the lining elements of the hot gas distributor should be carried out with as little gaps as possible between the individual lining elements in order to reduce the cooling air requirement, which is made possible by the radial bias.
- the transition lines and the hot gas distributor are preferably made of a highly heat-resistant metal material, in particular from a thin-walled, highly heat-resistant material in the manner of a sheet metal. The avoidance of nickel-based materials is a major advantage of the system described.
- the outer circumferential side and / or the inner circumferential side of the hot gas distributor is / are advantageously provided with a fastening flange which is laid out for fastening to the turbine. In this way, a very simple structure is achieved.
- the present invention creates a gas turbine plant with several annularly arranged burners, a turbine and a tubular combustion chamber system according to the invention, which connects the burner to the turbine.
- Figure 1 is a partially sectioned perspective
- FIG. 2 shows a perspective view of the arrangement shown in FIG. 1, viewed in the direction of arrow II in FIG.
- the figures show a tubular combustion chamber system 1 according to an embodiment of the present invention, which is connected to a turbine 2 of a gas turbine system 3.
- the tubular combustion chamber system 1 comprises a plurality of annularly arranged transition lines 4, which are designed with their upstream ends each to be connected to a burner 10 and to conduct hot gas generated by the burner 10 to the turbine 2, with only a single burner 10 being shown in FIG.
- the tubular combustion chamber system 1 comprises a hot gas distributor 5 designed to be connected to the turbine 2 and defining an annular channel 6 open to the turbine 2 into which the downstream ends of the transition lines 4 open.
- the transition lines 4 as well as the hot gas distributor 5 are made of metal, for example of a high-temperature metal alloy.
- the transition lines 4 each include a refractory lining 7, which is preferably made of a ceramic material.
- the transition lines 4 each have a cross section which tapers conically in the downstream direction.
- the refractory lining 7 of the transition lines 4 each comprises a plurality of ring-shaped lining sections with conically ver younger outer diameter in the downstream direction, which are formed in the present case by annular lining elements 7a.
- the lining elements 7a of a transition line 4 are pushed axially into the transition line 4, starting from the upstream end of the transition line 4, with spring and / or damping elements (not shown in detail) positioned along the circumference between the lining elements 7a and the inner wall of the transition line 4 which are guided in a form-fitting manner on the outer circumference of the lining elements 7a or on the inner wall of the transition line 4. Due to the conical design of the transition line 4 and the lining elements 7a, the lining elements 7a are radially and axially braced in such a way that they are held on the transition line 4 with radial and axial bias.
- the tension is maintained in the present case by an annular pressure element 8, which is inserted at the upstream end into the transition line 4, against the face of the adjacent lining element 7a is pressed and then attached to the transition line 4 while generating the desired pressing force. It can be attached, for example, by means of screws.
- the fireproof lining 7 of the hot gas distributor 5 is realized via a large number of lining elements 7b, which are advantageously also attached to the radially inside and outside surfaces of the hot gas distributor 5 under radial prestress.
- the outer peripheral side and the inner peripheral side of the hot gas distributor 5 on the free end of the hot gas distributor 5 facing the turbine 2 are provided with fastening flanges 9, which are designed for fastening to the turbine 2 by means of screws.
- the arrangement described above is advantageous in that, thanks to the additional hot gas distributor 5 of the tubular combustion chamber system 1 according to the invention, a very uniform flow of hot gas to the turbine 2 is achieved, whereby high-frequency temperature and speed changes are significantly reduced. This has a very positive effect on the service life of the turbine blades.
- the refractory lining 7 of the transition lines 4 and the hot gas distributor 5 is significantly reduced.
- the associated lower expansion differences in the area of the seals to the turbine 2 and the seals between the transition lines 4 lead to less wear in this area and enable more solid joining concepts between the tubular combustion chamber system 1 and the turbine 2.
- the refractory lining 7 pulls lower high temperature requirements for the materials of the metallic components 4 and 5, which means that costs can be saved.
- the transition lines 4 can be implemented without an inner layer system, which reduces the effort for repair and reprocessing is significantly reduced, since a stripping and recoating of the transition lines 4 is not necessary.
- the use of a refractory lining 7 also reduces the cooling requirement of the metallic components 4 and 5 of the tubular combustion chamber system 1. Compared to tubular combustion chamber systems without ceramic lining, the cooling air requirement can be reduced by up to 50% according to the calculations made earlier, which is an increase the performance of the gas turbine system 3 entails.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Combustion & Propulsion (AREA)
- Ceramic Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
Claims
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
DE102019204544.8A DE102019204544A1 (de) | 2019-04-01 | 2019-04-01 | Rohrbrennkammersystem und Gasturbinenanlage mit einem solchen Rohrbrennkammersystem |
PCT/EP2020/055501 WO2020200609A1 (de) | 2019-04-01 | 2020-03-03 | Rohrbrennkammersystem und gasturbinenanlage mit einem solchen rohrbrennkammersystem |
Publications (2)
Publication Number | Publication Date |
---|---|
EP3921577A1 true EP3921577A1 (de) | 2021-12-15 |
EP3921577B1 EP3921577B1 (de) | 2023-07-05 |
Family
ID=69810788
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP20711063.6A Active EP3921577B1 (de) | 2019-04-01 | 2020-03-03 | Rohrbrennkammersystem und gasturbinenanlage mit einem solchen rohrbrennkammersystem |
Country Status (4)
Country | Link |
---|---|
US (1) | US11852344B2 (de) |
EP (1) | EP3921577B1 (de) |
DE (1) | DE102019204544A1 (de) |
WO (1) | WO2020200609A1 (de) |
Families Citing this family (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE102019204544A1 (de) * | 2019-04-01 | 2020-10-01 | Siemens Aktiengesellschaft | Rohrbrennkammersystem und Gasturbinenanlage mit einem solchen Rohrbrennkammersystem |
Family Cites Families (41)
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GB626044A (en) * | 1945-06-21 | 1949-07-08 | Bristol Aeroplane Co Ltd | Improvements in or relating to gas turbine power plants |
US3981142A (en) * | 1974-04-01 | 1976-09-21 | General Motors Corporation | Ceramic combustion liner |
US4373326A (en) * | 1980-10-22 | 1983-02-15 | General Motors Corporation | Ceramic duct system for turbine engine |
DE3823510A1 (de) * | 1988-07-12 | 1990-01-18 | Kernforschungsanlage Juelich | Keramische auskleidung fuer einen brennraum |
GB2300909B (en) * | 1995-05-18 | 1998-09-30 | Europ Gas Turbines Ltd | A gas turbine gas duct arrangement |
JP3478531B2 (ja) * | 2000-04-21 | 2003-12-15 | 川崎重工業株式会社 | ガスタービンのセラミック部品支持構造 |
US6508052B1 (en) * | 2001-08-01 | 2003-01-21 | Rolls-Royce Corporation | Particle separator |
EP1528343A1 (de) * | 2003-10-27 | 2005-05-04 | Siemens Aktiengesellschaft | Keramischer Hitzeschildstein mit eingebetteten Verstärkungselementen zur Auskleidung einer Gasturbinenbrennkammerwand |
US7096668B2 (en) * | 2003-12-22 | 2006-08-29 | Martling Vincent C | Cooling and sealing design for a gas turbine combustion system |
EP1741980A1 (de) * | 2005-07-04 | 2007-01-10 | Siemens Aktiengesellschaft | Keramisches Bauteil mit heissgasresistenter Oberfläche und Verfahren zu seiner Herstellung |
US7908867B2 (en) * | 2007-09-14 | 2011-03-22 | Siemens Energy, Inc. | Wavy CMC wall hybrid ceramic apparatus |
US9127565B2 (en) * | 2008-04-16 | 2015-09-08 | Siemens Energy, Inc. | Apparatus comprising a CMC-comprising body and compliant porous element preloaded within an outer metal shell |
US8230688B2 (en) | 2008-09-29 | 2012-07-31 | Siemens Energy, Inc. | Modular transvane assembly |
US8402764B1 (en) * | 2009-09-21 | 2013-03-26 | Florida Turbine Technologies, Inc. | Transition duct with spiral cooling channels |
EP2309099B1 (de) * | 2009-09-30 | 2015-04-29 | Siemens Aktiengesellschaft | Verbindungskanal |
US9291063B2 (en) * | 2012-02-29 | 2016-03-22 | Siemens Energy, Inc. | Mid-section of a can-annular gas turbine engine with an improved rotation of air flow from the compressor to the turbine |
US20130239585A1 (en) * | 2012-03-14 | 2013-09-19 | Jay A. Morrison | Tangential flow duct with full annular exit component |
US9534497B2 (en) * | 2012-05-02 | 2017-01-03 | Honeywell International Inc. | Inter-turbine ducts with variable area ratios |
US9249678B2 (en) * | 2012-06-27 | 2016-02-02 | General Electric Company | Transition duct for a gas turbine |
RU2561956C2 (ru) * | 2012-07-09 | 2015-09-10 | Альстом Текнолоджи Лтд | Газотурбинная система сгорания |
US9309774B2 (en) * | 2014-01-15 | 2016-04-12 | Siemens Energy, Inc. | Assembly for directing combustion gas |
WO2015199693A1 (en) * | 2014-06-26 | 2015-12-30 | Siemens Energy, Inc. | Converging flow joint insert system at an intersection between adjacent transition duct bodljs |
US9702258B2 (en) * | 2014-07-01 | 2017-07-11 | Siemens Energy, Inc. | Adjustable transition support and method of using the same |
US10024180B2 (en) * | 2014-11-20 | 2018-07-17 | Siemens Energy, Inc. | Transition duct arrangement in a gas turbine engine |
US11906079B2 (en) * | 2015-04-16 | 2024-02-20 | Krzysztof Jan Wajnikonis | Telescopically assembled mechanical connector |
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EP3301374A1 (de) * | 2016-09-29 | 2018-04-04 | Siemens Aktiengesellschaft | Pilotbrenneranordnung mit pilotluftversorgung |
US20180106155A1 (en) * | 2016-10-13 | 2018-04-19 | Siemens Energy, Inc. | Transition duct formed of a plurality of segments |
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EP3486431B1 (de) * | 2017-11-15 | 2023-01-04 | Ansaldo Energia Switzerland AG | Heissgaspfadkomponente für ein gasturbinentriebwerk sowie gasturbinentriebwerk mit derselben |
EP3499125A1 (de) | 2017-12-12 | 2019-06-19 | Siemens Aktiengesellschaft | Rohrbrennkammer mit keramischer auskleidung |
WO2020086069A1 (en) * | 2018-10-24 | 2020-04-30 | Siemens Energy, Inc. | Transition duct system with non-metallic thermally-insulating liners supported with splittable metallic shell structures for delivering hot-temperature gasses in a combustion turbine engine |
JP7149807B2 (ja) * | 2018-11-01 | 2022-10-07 | 三菱重工業株式会社 | ガスタービン燃焼器 |
GB201902693D0 (en) * | 2019-02-28 | 2019-04-17 | Rolls Royce Plc | Combustion liner and gas turbine engine comprising a combustion liner |
DE102019204544A1 (de) * | 2019-04-01 | 2020-10-01 | Siemens Aktiengesellschaft | Rohrbrennkammersystem und Gasturbinenanlage mit einem solchen Rohrbrennkammersystem |
US11215367B2 (en) * | 2019-10-03 | 2022-01-04 | Raytheon Technologies Corporation | Mounting a ceramic component to a non-ceramic component in a gas turbine engine |
-
2019
- 2019-04-01 DE DE102019204544.8A patent/DE102019204544A1/de not_active Withdrawn
-
2020
- 2020-03-03 EP EP20711063.6A patent/EP3921577B1/de active Active
- 2020-03-03 WO PCT/EP2020/055501 patent/WO2020200609A1/de unknown
- 2020-03-03 US US17/440,354 patent/US11852344B2/en active Active
Also Published As
Publication number | Publication date |
---|---|
US11852344B2 (en) | 2023-12-26 |
WO2020200609A1 (de) | 2020-10-08 |
DE102019204544A1 (de) | 2020-10-01 |
US20220186928A1 (en) | 2022-06-16 |
EP3921577B1 (de) | 2023-07-05 |
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