EP3318803B1 - Bolzenanordnung für gasturbinenbrennkammer - Google Patents
Bolzenanordnung für gasturbinenbrennkammer Download PDFInfo
- Publication number
- EP3318803B1 EP3318803B1 EP17200015.0A EP17200015A EP3318803B1 EP 3318803 B1 EP3318803 B1 EP 3318803B1 EP 17200015 A EP17200015 A EP 17200015A EP 3318803 B1 EP3318803 B1 EP 3318803B1
- Authority
- EP
- European Patent Office
- Prior art keywords
- combustor
- aft
- free zone
- liner panel
- recited
- 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.)
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- 238000001816 cooling Methods 0.000 claims description 14
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- 238000002485 combustion reaction Methods 0.000 description 12
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Images
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
-
- 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
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00017—Assembling combustion chamber liners or subparts
-
- 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
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03042—Film cooled combustion chamber walls or domes
Definitions
- the present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
- Gas turbine engines such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
- the combustor section typically includes a combustion chamber formed by an inner and outer wall assembly.
- Each wall assembly includes a support shell lined with heat shields often referred to as liner panels.
- combustor Impingement Film-Cooled Floatwall (IFF) liner panels are typically a curved flat surface on a hot side exposed to the gas path.
- the opposite, or cold side has features such as cast in threaded studs to mount the liner panel and a full perimeter rail that contacts the inner surface of the respective liner shell. These features may result in durability issues.
- US 2015/362 192 A1 discloses a prior art liner panel for use in a combustor of a gas turbine engine as set forth in the preamble of claim 1.
- US 2010/095 680 A1 discloses a prior art dual wall structure for use in a combustor of a gas turbine engine.
- US 2010/095679 A1 disclose a prior art dual wall structure for use as a combustor of a gas turbine engine.
- the invention provides a combustor for a gas turbine engine as recited in claim 1.
- the invention provides a method of directing airflow through a wall assembly within a combustor of a gas turbine engine as recited in claim 9.
- FIG. 1 schematically illustrates a gas turbine engine 20.
- the gas turbine engine 20 is disclosed herein as a two-spool turbo fan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
- Alternative engine architectures might include an augmentor section among other systems or features.
- the fan section 22 drives air along a bypass flowpath and into the compressor section 24.
- the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26, which then expands and directs the air through the turbine section 28.
- turbofan in the disclosed non-limiting embodiment, it should be appreciated that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an Intermediate Pressure Compressor ("IPC") between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an Intermediate Pressure Turbine (“IPT”) between the High Pressure Turbine (“HPT”) and the Low Pressure Turbine (“LPT”).
- IPC Intermediate Pressure Compressor
- LPC Low Pressure Compressor
- HPC High Pressure Compressor
- IPT Intermediate Pressure Turbine
- the engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing structures 38.
- the low spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a Low Pressure Compressor (“LPC”) 44 and a Low Pressure Turbine (“LPT”) 46.
- the inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30.
- An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.
- the high spool 32 includes an outer shaft 50 that interconnects a High Pressure Compressor ("HPC") 52 and High Pressure Turbine (“HPT”) 54.
- a combustor 56 is arranged between the HPC 52 and the HPT 54.
- the inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- Core airflow is compressed by the LPC 44, then the HPC 52, mixed with the fuel and burned in the combustor 56, then expanded over the HPT 54 and the LPT 46.
- the LPT 46 and HPT 54 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion.
- the main engine shafts 40, 50 are supported at a plurality of points by bearing systems 38 within the static structure 36.
- the gas turbine engine 20 is a high-bypass geared aircraft engine.
- the gas turbine engine 20 bypass ratio is greater than about six.
- the geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system.
- the example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1.
- the geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the LPC 44 and LPT 46 and render increased pressure in a fewer number of stages.
- a pressure ratio associated with the LPT 46 is pressure measured prior to the inlet of the LPT 46 as related to the pressure at the outlet of the LPT 46 prior to an exhaust nozzle of the gas turbine engine 20.
- the bypass ratio of the gas turbine engine 20 is greater than about ten
- the fan diameter is significantly larger than that of the LPC 44
- the LPT 46 has a pressure ratio that is greater than about five. It should be appreciated, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
- a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio.
- the fan section 22 of the gas turbine engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet (10,668m). This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC).
- TSFC Thrust Specific Fuel Consumption
- Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system.
- the low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine 20 is less than 1.45.
- the Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
- the combustor section 26 generally includes a combustor 56 with an outer combustor wall assembly 60, an inner combustor wall assembly 62, and a diffuser case module 64.
- the outer combustor wall assembly 60 and the inner combustor wall assembly 62 are spaced apart such that a combustion chamber 66 is defined therebetween.
- the combustion chamber 66 is generally annular in shape to surround the engine central longitudinal axis A.
- the outer combustor liner assembly 60 is spaced radially inward from an outer diffuser case 64A of the diffuser case module 64 to define an outer annular plenum 76.
- the inner combustor liner assembly 62 is spaced radially outward from an inner diffuser case 64B of the diffuser case module 64 to define an inner annular plenum 78. It should be appreciated that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further appreciated that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.
- the combustor wall assemblies 60, 62 contain the combustion products for direction toward the turbine section 28.
- Each combustor wall assembly 60, 62 generally includes a respective support shell 68, 70 which supports one or more liner panels 72, 74 mounted thereto arranged to form a liner array.
- the support shells 68, 70 may be manufactured by, for example, the hydroforming of a sheet metal alloy to provide the generally cylindrical outer shell 68 and inner shell 70.
- Each of the liner panels 72, 74 may be generally rectilinear with a circumferential arc.
- the liner panels 72, 74 may be manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material.
- the liner array includes a multiple of forward liner panels 72A and a multiple of aft liner panels 72B that are circumferentially staggered to line the outer shell 68.
- a multiple of forward liner panels 74A and a multiple of aft liner panels 74B are circumferentially staggered to line the inner shell 70.
- the combustor 56 further includes a forward assembly 80 immediately downstream of the compressor section 24 to receive compressed airflow therefrom.
- the forward assembly 80 generally includes a cowl 82, a bulkhead assembly 84, and a multiple of swirlers 90 (one shown). Each of the swirlers 90 is circumferentially aligned with one of a multiple of fuel nozzles 86 (one shown) and the respective hood ports 94 to project through the bulkhead assembly 84.
- the bulkhead assembly 84 includes a bulkhead support shell 96 secured to the combustor walls 60, 62, and a multiple of circumferentially distributed bulkhead liner panels 98 secured to the bulkhead support shell 96 around the swirler opening.
- the bulkhead support shell 96 is generally annular and the multiple of circumferentially distributed bulkhead liner panels 98 are segmented, typically one to each fuel nozzle 86 and swirler 90.
- the cowl 82 extends radially between, and is secured to, the forwardmost ends of the combustor walls 60, 62.
- the cowl 82 includes a multiple of circumferentially distributed hood ports 94 that receive one of the respective multiple of fuel nozzles 86 and facilitates the direction of compressed air into the forward end of the combustion chamber 66 through a swirler opening 92.
- Each fuel nozzle 86 may be secured to the diffuser case module 64 and project through one of the hood ports 94 and through the swirler opening 92 within the respective swirler 90.
- the forward assembly 80 introduces core combustion air into the forward section of the combustion chamber 66 while the remainder enters the outer annular plenum 76 and the inner annular plenum 78.
- the multiple of fuel nozzles 86 and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber 66.
- the outer and inner support shells 68, 70 are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in the HPT 54.
- the NGVs 54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section 28 to facilitate the conversion of pressure energy into kinetic energy.
- the core airflow combustion gases are also accelerated by the NGVs 54A because of their convergent shape and are typically given a "spin” or a "swirl” in the direction of turbine rotor rotation.
- the turbine rotor blades absorb this energy to drive the turbine rotor at high speed.
- a multiple of studs 100 extend from each of the liner panels 72, 74 so as to permit a liner array (partially shown in Figure 4 ) of the liner panels 72, 74 to be mounted to their respective support shells 68, 70 with fasteners 102 such as nuts. That is, the studs 100 project rigidly from the liner panels 72, 74 to extend through the respective support shells 68, 70 and receive the fasteners 102 on a threaded section thereof ( Figure 5 ).
- a multiple of cooling impingement passages 104 penetrate through the support shells 68, 70 to allow air from the respective annular plenums 76, 78 to enter cavities 106 formed in the combustor walls 60, 62 between the respective support shells 68, 70 and liner panels 72, 74.
- the impingement passages 104 are generally normal to the surface of the liner panels 72, 74.
- the air in the cavities 106 provides cold side impingement cooling of the liner panels 72, 74 that is generally defined herein as heat removal via internal convection.
- a multiple of effusion passages 108 penetrate through each of the liner panels 72, 74.
- the geometry of the passages e.g., diameter, shape, density, surface angle, incidence angle, etc., as well as the location of the passages with respect to the high temperature combustion flow also contributes to effusion cooling.
- the effusion passages 108 allow the air to pass from the cavities 106 defined in part by a cold side 110 of the liner panels 72, 74 to a hot side 112 of the liner panels 72, 74 and thereby facilitate the formation of a thin, relatively cool, film of cooling air along the hot side 112.
- each of the multiple of effusion passages 108 are typically 0.025" (0.635 mm) in diameter and define a surface angle of about thirty (30) degrees with respect to the cold side 110 of the liner panels 72, 74.
- the effusion passages 108 are generally more numerous than the impingement passages 104 and promote film cooling along the hot side 112 to sheath the liner panels 72, 74 ( Figure 6 ).
- Film cooling as defined herein is the introduction of a relatively cooler air at one or more discrete locations along a surface exposed to a high temperature environment to protect that surface in the region of the air injection as well as downstream thereof.
- impingement passages 104 and effusion passages 108 may be referred to as an Impingement Film Floatwall (IFF) assembly.
- IFF Impingement Film Floatwall
- a multiple of dilution passages 116 are located in the liner panels 72, 74 each along a common axis D.
- the dilution passages 116 are located in a circumferential line W (shown partially in Figure 4 ).
- the dilution passages 116 are illustrated in the disclosed non-limiting embodiment as within the aft liner panels 72B, 74B, the dilution passages may alternatively be located in the forward liner panels 72A, 72B or in a single liner panel which replaces the fore/aft liner panel array.
- the dilution passages 116 although illustrated in the disclosed non-limiting embodiment as integrally formed in the liner panels, it should be appreciated that the dilution passages 116 may be separate components. Whether integrally formed or separate components, the dilution passages 116 may be referred to as grommets.
- each of the liner panels 72A, 72B, 74A, 74B in the liner panel array includes a perimeter rail 120 formed by a forward circumferential rail 122, an aft circumferential rail 124, and axial rails 126A, 126B, that interconnect the forward and aft circumferential rail 122, 124.
- the perimeter rail 120 seals each liner panel with respect to the respective support shell 68, 70 to form the impingement cavity 106 therebetween.
- the forward and aft circumferential rail 122, 124 are located at relatively constant curvature shell interfaces while the axial rails 126 extend across an axial length of the respective support shell 68, 70 to complete the perimeter rail 120 that seals the liner panels 72, 74 to the respective support shell 68, 70.
- a multiple of studs 100 are located adjacent to the respective forward circumferential rail 122 and the aft circumferential rail 124.
- Each of the studs 100 may be at least partially surrounded by posts 130 to at least partially support the fastener 102 and provide a stand-off between each liner panels 72B, 74B and respective support shell 68, 70.
- the quantity and location of the multiple of studs 100 is typically based on structural analysis and symmetry of the studs 100 relative to the liner to facilitate proper sealing of the panel rail to the inner combustor shell.
- the conventional position would often locate one or more of the multiple of studs 100 downstream of the combustor swirlers 90. As this area may have relatively high metal temperatures, durability issues may result from the lack of effusion cooling as this issue may be more significant at the aft section of the forward row of liners.
- the multiple of studs 100 of the forward liner panels 72A, 74A are not located within a stud free zone 200 defined downstream of each of the combustion swirlers 90.
- Each stud free zone 200 is defined as an essentially truncated triangular shape.
- the stud free zone 200 is defined by a truncated triangle with a truncated apex 201 located at the combustor swirler 90.
- the stud free zone 200 is a trapezoidal shaped zone located at the aft edge of the forward liner panels 72A, 74A. That is, to increase durability, the studs 100 are specifically moved away from each zone 200 directly aft of the respective combustor swirlers 90 in the aft section of the forward liner panels 72A, 74A as this area has the relatively hottest surface metal temperatures.
- the stud free zone 200 facilitates a more efficient distribution of film cooling holes in the hottest areas of the segment as the studs no longer hinder location of film cooling holes 108 ( Figure 3 ).
- the stud free zone 200 extends from about 0.75 inches (19 mm) to 1.7 inches (43 mm) from the respective combustor swirler 90 and the sides between the fore and aft lines are at about 20 degrees.
- the forward most row of studs 100A are also intentionally moved out away from an aft liner panel stud free zone 202 downstream of the stud free zone 200 and the respective combustor swirler 90. As with the forward liner panels 72A, 74A, this permits a more advantageous distribution of cooling holes around the area of the liner segments which typically have the hottest metal temperatures. Further, at least one dilution passage 116 may be located within the aft liner panel stud free zone 202.
- the stud free zone 202 in the aft liner panels 72B, 74B defines a rectangle of about 1.5 inches (38 mm) by 2.8 inches (71mm) and is located 1.8 inches (45mm) behind the respective combustor swirlers 90.
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Claims (10)
- Brennkammer (56) für ein Gasturbinentriebwerk (20), die Folgendes umfasst:eine Vielzahl von Brennkammerverwirblern (90);eine Stützhülle (68, 70); undeine vordere Auskleidungsplatte (72A, 74A), die über eine Vielzahl von Bolzen (100) an der Stützhülle (68, 70) montiert ist, wobei die vordere Auskleidungsplatte (72A, 74A) eine bolzenfreie Zone (200) stromabwärts jedes entsprechenden Brennkammerverwirblers (90) beinhaltet, dadurch gekennzeichnet dass:
die bolzenfreite Zone (200) trapezförmig ist, sich zu einer hinteren Kante der vorderen Auskleidungsplatte (72A, 74A hin befindet und durch ein verkürztes Dreieck mit einem verkürzten Scheitelpunkt (201) definiert ist, der sich an dem Brennkammerverwirbler (90) befindet. - Brennkammer (56) nach Anspruch 1, die ferner eine hintere Auskleidungsplatte (72B, 74B) hinter der vorderen Auskleidungsplatte (72A, 74A) umfasst.
- Brennkammer (56) nach Anspruch 2, die ferner eine hintere bolzenfreie Zone (202) stromabwärts der vorderen bolzenfreien Zone (200) der Auskleidungsplatte umfasst.
- Brennkammer (56) nach Anspruch 3, die ferner mindestens eine Hauptdiffusionsöffnung (116) in der hinteren bolzenfreien Zone (202) stromabwärts der vorderen bolzenfreien Zone (200) der vorderen Auskleidungsplatte umfasst.
- Brennkammer (56) nach einem der vorhergehenden Ansprüche, wobei die bolzenfreie Zone (200) ein Vielfaches von Filmkühllöchern (108) beinhaltet.
- Brennkammer (56) nach einem der vorhergehenden Ansprüche, die ferner eine vordere Baugruppe (80) umfasst, die eine Trennwandstützhülle (96), eine Trennwandbaugruppe (84), die an der Trennwandstützhülle (96) montiert ist, und ein Vielfaches der Brennkammerverwirbler (90) beinhaltet, die mindestens teilweise durch die Trennwandbaugruppe (84) montiert sind.
- Brennkammer (56) nach Anspruch 6, wobei die vordere Baugruppe (80) an der Stützhülle (68, 70) montiert ist.
- Brennkammer (56) nach Anspruch 6 oder 7, die ferner ein Vielfaches von in Umfangsrichtung verteilten Trennwandauskleidungsplatten (98) umfasst, die an die Trennwandstützhülle (96) um eine Verwirbleröffnung (92) gesichert sind.
- Verfahren zum Leiten eines Luftstroms durch eine Wandbaugruppe (60, 62) innerhalb einer Brennkammer (56) eines Gasturbinentriebwerks (20), das ein Bereitstellen einer bolzenfreien Zone (200) in einer vorderen Auskleidungsplatte (72A, 74A) stromabwärts eines entsprechenden Brennkammerverwirblers (90) umfasst, wobei die bolzenfreie Zone (200) ein Vielfaches von Filmkühllöchern (108) beinhaltet, wobei die bolzenfreie Zone (200) trapezförmig ist, sich zu einer hinteren Kante der vorderen Auskleidungsplatte (72A, 74A) hin befindet und durch ein verkürztes Dreieck mit einem verkürzten Scheitelpunkt (201) definiert ist, der sich an dem Brennkammerverwirbler (90) befindet.
- Verfahren nach Anspruch 9, das ferner ein Lokalisieren eines Verdünnungskanals (116) innerhalb einer hinteren bolzenfreien Zone (202) in einer hinteren Auskleidungsplatte (72B, 74B) umfasst, wobei die hintere Auskleidungsplatte (72B, 74B) stromabwärts der vorderen Auskleidungsplatte (72A, 74A) liegt.
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US15/343,988 US20180128485A1 (en) | 2016-11-04 | 2016-11-04 | Stud arrangement for gas turbine engine combustor |
Publications (2)
Publication Number | Publication Date |
---|---|
EP3318803A1 EP3318803A1 (de) | 2018-05-09 |
EP3318803B1 true EP3318803B1 (de) | 2021-05-05 |
Family
ID=60245009
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP17200015.0A Active EP3318803B1 (de) | 2016-11-04 | 2017-11-03 | Bolzenanordnung für gasturbinenbrennkammer |
Country Status (2)
Country | Link |
---|---|
US (1) | US20180128485A1 (de) |
EP (1) | EP3318803B1 (de) |
Family Cites Families (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4614082A (en) * | 1972-12-19 | 1986-09-30 | General Electric Company | Combustion chamber construction |
GB2298267B (en) * | 1995-02-23 | 1999-01-13 | Rolls Royce Plc | An arrangement of heat resistant tiles for a gas turbine engine combustor |
US5782294A (en) * | 1995-12-18 | 1998-07-21 | United Technologies Corporation | Cooled liner apparatus |
US7509809B2 (en) * | 2005-06-10 | 2009-03-31 | Pratt & Whitney Canada Corp. | Gas turbine engine combustor with improved cooling |
US20100095679A1 (en) * | 2008-10-22 | 2010-04-22 | Honeywell International Inc. | Dual wall structure for use in a combustor of a gas turbine engine |
US20100095680A1 (en) * | 2008-10-22 | 2010-04-22 | Honeywell International Inc. | Dual wall structure for use in a combustor of a gas turbine engine |
US20150260399A1 (en) * | 2012-09-28 | 2015-09-17 | United Technologies Corporation | Combustor section of a gas turbine engine |
WO2014052965A1 (en) * | 2012-09-30 | 2014-04-03 | United Technologies Corporation | Interface heat shield for a combustor of a gas turbine engine |
WO2014113007A1 (en) * | 2013-01-17 | 2014-07-24 | United Technologies Corporation | Gas turbine engine combustor liner assembly with convergent hyperbolic profile |
US8976451B2 (en) * | 2013-08-02 | 2015-03-10 | Forward Optics Co., Ltd | Lens array module |
-
2016
- 2016-11-04 US US15/343,988 patent/US20180128485A1/en not_active Abandoned
-
2017
- 2017-11-03 EP EP17200015.0A patent/EP3318803B1/de active Active
Also Published As
Publication number | Publication date |
---|---|
US20180128485A1 (en) | 2018-05-10 |
EP3318803A1 (de) | 2018-05-09 |
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