US20180128487A1 - Combustion chamber of a gas turbine - Google Patents

Combustion chamber of a gas turbine Download PDF

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Publication number
US20180128487A1
US20180128487A1 US15/808,162 US201715808162A US2018128487A1 US 20180128487 A1 US20180128487 A1 US 20180128487A1 US 201715808162 A US201715808162 A US 201715808162A US 2018128487 A1 US2018128487 A1 US 2018128487A1
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US
United States
Prior art keywords
shingle
cooling holes
combustion chamber
distance
arrangement
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.)
Abandoned
Application number
US15/808,162
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English (en)
Inventor
Miklos Gerendas
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Rolls Royce Deutschland Ltd and Co KG filed Critical Rolls Royce Deutschland Ltd and Co KG
Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: GERENDAS, MIKLOS
Publication of US20180128487A1 publication Critical patent/US20180128487A1/en
Abandoned legal-status Critical Current

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    • 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
    • F23R3/04—Air inlet arrangements
    • 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
    • F23R2900/03041—Effusion cooled combustion chamber walls or domes
    • 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
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies
    • 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

Definitions

  • the invention relates to a combustion chamber of a gas turbine according to the features of the generic term of claim 1 .
  • the invention relates to a combustion chamber of a gas turbine which is covered with shingles.
  • At least one shingle comprises a plate-shaped shingle body which has a circumferential shingle edge. It extends from the cold side of the shingle body that is facing away from the combustion chamber's interior space to the combustion chamber wall and thus forms an intermediate space between the shingle body and the combustion chamber wall. Cooling air is introduced into this intermediate space through impingement cooling holes, and is subsequently discharged through effusion cooling holes of the shingle body located at its surface.
  • the shingle body is provided with an arrangement of effusion cooling holes, which may for example be embodied in a row-shaped manner or with another arrangement with respect to each other.
  • the impingement cooling holes of the shingle wall are also embodied in a suitable arrangement.
  • EP 0 576 435 B1 shows a structure that is illustrated in FIG. 2 .
  • a shingle 25 has a substantially plate-shaped shingle body 29 that is delimited by a shingle edge 31 .
  • the shingle edge 31 extends from the shingle body 29 in the direction towards the combustion chamber wall 32 to form an intermediate space 35 .
  • the combustion chamber wall 32 is provided with impingement cooling holes 34 to introduce cooling air into the intermediate space 35 . This cooling air flows out of the intermediate space 35 through the effusion cooling holes 33 that are formed in the shingle body 29 .
  • leakage air 36 Due to the fact that the shingle edge 31 cannot be arranged in a sealing manner at the combustion chamber wall 32 , there is always a leakage, which is illustrated as leakage air 36 . Thus, a part of the air volume that is supplied through the impingement cooling holes 34 flows from the intermediate space 35 unused as leakage air 36 , and cannot be used to flow through the effusion cooling holes 33 .
  • FIG. 3 shows a similar structure, wherein the same parts are indicated by the same reference signs.
  • U.S. Pat. No. 5,598,697 A shows that either leakage air 36 flows out of the intermediate space 35 and cannot be used for effusion cooling holes 33 , or a seal 38 has to be used.
  • EP 1 351 022 B1 is quoted as the state of the art.
  • the invention is based on the objective of creating a combustion chamber of a gas turbine which ensures an effective use of the cooling air in the area of the shingle, while at the same time having a simple structure and a single, cost-effective manufacturability.
  • the arrangement of impingement cooling holes has a distance to the shingle edge which lies between 1.5 to 2 times the distance of the surface of the combustion chamber wall to the surface of the shingle body, and that in the corner areas of the shingle the distance of the arrangement of impingement cooling holes is between 1.1 and 3 times the above-mentioned distance.
  • the invention is based on the basic principle of designing the inflow of cooling air into the intermediate space formed by the shingle in such a manner that the supply of cooling air through the impingement cooling holes occurs in the middle area of the shingle body, i.e. up to a distance from the shingle edge.
  • the cooling air can be discharged through the sufficiently dimensioned effusion cooling holes. This flow of air is caused by the resulting pressure difference across the shingle. In a completely sealed shingle edge, the air flows though all of the effusion bore holes, without any dead bands of the flow being formed.
  • the solution according to the invention makes it possible to use substantially the entire volume of cooling air for the purpose of cooling the shingles, namely, on the one hand, for cooling the cold surface of the shingle that is facing away from the combustion chamber interior space by means of impingement cooling and, on the other hand, for film cooling by means of the air that is discharged through the effusion cooling holes. Since what results according to the invention is a considerable or complete reduction of the leakage flow, the present invention results in a considerable increase of the efficiency of the shingle cooling.
  • the impingement cooling holes are not formed up to the shingle edge, but that the arrangement of impingement cooling holes is chosen in such a manner that each impingement cooling hole has a distance from each shingle edge through which a leakage may occur. This distance is chosen in such a manner that it is defined based on the free jet length of the impingement cooling jet.
  • the free jet length is the path length between the exit site of the cooling air from the impingement cooling hole and impingement site on the cold surface of the shingle body that is facing away from the combustion chamber's interior space.
  • the volume of the intermediate space between the combustion chamber wall and the shingle body is also defined based on this free jet length.
  • the distance of the impingement cooling holes from the shingle edge is dimensioned in such a manner that it corresponds to at least 1.5 times the free jet length of the impingement cooling jet.
  • the distance can be up to 2 times the free jet length, with this value being a preferred value.
  • Combustion chamber shingles are usually formed in a rectangular, more seldom in an triangular or diamond-shaped, manner.
  • the result is a corner area of the shingle edges in the two neighboring edge areas, which meet in the corner area and in which no impingement cooling holes are present, meet.
  • the edge distances between the shingle edge and the arrangement of impingement cooling holes are linearly added.
  • a bevel or rounding is formed here in the arrangement of the impingement cooling holes.
  • this beveled or rounded area is defined by a factor that can be referred to as the overlay constant.
  • This overlay constant has a value of 1.1 to 3, preferably of 1.5 to 2.5, ideally of 2.
  • the distance in the corner area is enlarged by the factor of the overlay constant to ensure that the intermediate space between the shingle body and the combustion chamber wall is passed by the flow to the desired extent.
  • a smaller projected surface results for the arrangement of impingement cooling holes than for the arrangement of effusion cooling holes.
  • the arrangement of impingement cooling holes is thus shifted away from the shingle edge, and is set at a distance to the same.
  • the impingement cooling holes can be set closer to each other in the area of the arrangement of impingement cooling holes to avoid any impingement cooling holes located close to the edge within the distance to the shingle edge.
  • FIG. 1 shows a schematic rendering of a gas turbine engine according to the present invention
  • FIG. 2 shows a rendering of the state of the art
  • FIG. 3 shows a rendering of the state of the art
  • FIG. 4 shows a schematic side view of a first exemplary embodiment of the invention
  • FIG. 5 shows a rendering of a further exemplary embodiment of the invention, which is analogous to FIG. 4 .
  • FIG. 6 shows a simplified rendering, which is analogous to FIG. 5 , including a rendering of possible leakage flows,
  • FIG. 7 shows a top view of a first exemplary embodiment of the corner design
  • FIG. 8 shows a view of a further exemplary embodiment, which is analogous to FIG. 7 .
  • the gas turbine engine 10 represents a general example of a turbomachine in which the invention may be used.
  • the engine 10 is configured in a conventional manner and comprises, arranged successively in flow direction, an air intake 11 , a fan 12 that rotates inside a housing, a medium-pressure compressor 13 , a high-pressure compressor 14 , a combustion chamber 15 , a high-pressure turbine 16 , a medium-pressure turbine 17 , and a low-pressure turbine 18 as well as an exhaust nozzle 19 , which are all arranged around a central engine axis 1 .
  • the medium-pressure compressor 13 and the high-pressure compressor 114 respectively comprise multiple stages, of which each has an arrangement of fixedly arranged stationary guide vanes 20 that extends in the circumferential direction, with the stationary guide vanes 20 being generally referred to as stator vanes and projecting radially inward from the core engine shroud 21 through the compressors 13 , 14 into a ring-shaped flow channel.
  • the compressors have an arrangement of compressor rotor blades 22 that project radially outward from a rotatable drum or disc 26 , and are coupled to hubs 27 of the high-pressure turbine 16 or the medium-pressure turbine 17 .
  • the turbine sections 16 , 17 , 18 have similar stages, comprising an arrangement of stationary guide vanes 23 projecting radially inward from the housing 21 through the turbines 16 , 17 , 18 into the ring-shaped flow channel, and a subsequent arrangement of turbine blades/vanes 24 projecting outwards from the rotatable hub 27 .
  • the compressor drum or compressor disc 26 and the blades 22 arranged thereon as well as the turbine rotor hub 27 and the turbine rotor blades/vanes 24 arranged thereon rotate around the engine central axis 1 .
  • FIGS. 4 to 6 respectively show simplified sectional views in a sectional plane that comprises the central axis of a combustion chamber 15 , which is not shown.
  • a combustion chamber wall 32 provided with an arrangement of impingement cooling holes 34 is shown in a schematic manner.
  • the impingement cooling holes 34 as well as the effusion cooling holes 33 are shown only by the flow direction in the form of a flow arrow.
  • Shingles 25 are arranged at a side of the combustion chamber wall 32 that is facing towards the combustion chamber interior space 30 , being for example screwed on, as it is shown in FIG. 2 .
  • the shingles have a plate-shaped, substantially flat shingle body 29 that is provided with effusion cooling holes 33 .
  • a circumferential shingle edge 31 is formed, abutting the combustion chamber wall 32 .
  • the height of the shingle edge 31 defines the volume of an intermediate space 35 into which the impingement cooling air flows and is subsequently discharged through the effusion cooling holes 33 .
  • the arrangement of impingement cooling holes 34 is arranged at a distance A from the shingle edge 31 .
  • the effusion cooling holes 33 are distributed about the entire surface of the shingle body 29 .
  • FIG. 4 shows an exemplary embodiment in which the ratio of the number of impingement cooling holes to the effusion cooling holes is 1:1.
  • the distance A is chosen in such a manner in this exemplary embodiment that a row of effusion cooling holes is located between the edge of the next impingement cooling hole 34 and the shingle edge 31 , as shown in the right-hand half of FIG. 4 .
  • the ratio of impingement cooling holes to the effusion cooling holes is 1:2. Consequently, two rows of effusion cooling holes 33 are provided in the distance area A between the shingle edge 31 and the arrangement of impingement cooling holes 34 .
  • FIG. 6 shows a rendering that is analogous to FIG. 5 and from which it can be seen that, in the most unfavorable case, only a very small leakage air flow 36 would flow via the shingle edge 31 from the intermediate space 35 should the shingle edge 31 be sealed very insufficiently against the combustion chamber wall 32 .
  • FIGS. 7 and 8 respectively show a simplified top view of the embodiment according to the invention in a schematic rendering.
  • the shingle edge 31 is shown, which provides a seating surface of the shingle, as shown in FIGS. 4 to 6 .
  • a field of impingement cooling holes is indicated by the reference sign 37 , without describing the individual impingement cooling holes and their arrangement. They can be arranged in a suitable manner, with the particular arrangement of impingement cooling holes not playing a decisive role for the invention. Rather, what is important here is that a distance A, in which no impingement cooling holes and thus no impingement perforation is present, results between the side of the shingle edge 31 that is facing towards the arrangement of impingement cooling holes 34 .
  • FIGS. 7 and 8 show the inner side of the shingle edge 31 as edge R 1 or R 2 . Further, FIGS. 7 and 8 respectively show the distance A between the edge R 1 or R 2 and a boundary G of the field 37 of the impingement cooling holes.
  • the distances A add up at the edges of the field 37 according to the invention, resulting in a beveling of the field 37 .
  • the edge distances A add up in such a manner in the edges of the shingle 25 , that a value A results if a distance from the first edge R 1 of the seating surface of the shingle edge 31 of the shingle 25 .
  • a value A also results from the second edge R 2 of the seating surface of the shingle edge 31 of the shingle.
  • the field 37 of the impingement cooling holes ends along a line L 1 that is parallel to the edge R 1 , and at a distance along a line L 2 that is parallel to the edge R 2 .
  • the distance A is defined.
  • the boundary of the field 37 of the impingement cooling holes is indicated by G as a dashed line.
  • the field 37 of the impingement cooling pattern has a distance of C ⁇ A on the line L 1 along the edge R 1 .
  • a distance of C ⁇ A results regarding the edge R 2 and the line of the boundary G of the impingement cooling pattern.
  • FIG. 7 shows an ideal state with a distance of 2 ⁇ A from the impingement cooling perforation along the line L 1 to the edge R 2 .
  • the result is additional corner area in the shape of an equilateral triangle, with no impingement cooling holes being provided therein.
  • FIG. 8 shows a variant in which the field 37 of the impingement cooling holes is rounded off in the corner area, so that the boundary G extends in the form of a circular arc in this location.
  • the openings for the impingement cooling are thus placed at a distance A from the edge of the shingle in order to avoid any edge leakages from an impingement-effusion-cooled shingle, so that effusion cooling holes can be arranged between the impingement cooling opening that is closest to the edge and the inner side of the edge of the shingle so as to ensure an outflow of cooling air from the impingement cooling holes through the effusion cooling holes, and to avoid any edge leakage.
  • the distance from the edge of the shingle, in which no impingement cooling holes are provided is at least 2 times the free path length of the impingement cooling jet within the intermediate space formed by the shingle.

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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)
US15/808,162 2016-11-10 2017-11-09 Combustion chamber of a gas turbine Abandoned US20180128487A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102016222099.3A DE102016222099A1 (de) 2016-11-10 2016-11-10 Brennkammer einer Gasturbine
DE102016222099.3 2016-11-10

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US20180128487A1 true US20180128487A1 (en) 2018-05-10

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US15/808,162 Abandoned US20180128487A1 (en) 2016-11-10 2017-11-09 Combustion chamber of a gas turbine

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US (1) US20180128487A1 (fr)
EP (1) EP3321583B1 (fr)
DE (1) DE102016222099A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180299126A1 (en) * 2017-04-18 2018-10-18 United Technologies Corporation Combustor liner panel end rail
US20180306113A1 (en) * 2017-04-19 2018-10-25 United Technologies Corporation Combustor liner panel end rail matching heat transfer features

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102020203017A1 (de) 2020-03-10 2021-09-16 Siemens Aktiengesellschaft Brennkammer mit keramischem Hitzeschild und Dichtung

Citations (3)

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Publication number Priority date Publication date Assignee Title
US6701714B2 (en) * 2001-12-05 2004-03-09 United Technologies Corporation Gas turbine combustor
US7219498B2 (en) * 2004-09-10 2007-05-22 Honeywell International, Inc. Waffled impingement effusion method
US7363763B2 (en) * 2003-10-23 2008-04-29 United Technologies Corporation Combustor

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GB2160964B (en) * 1984-06-25 1988-04-07 Gen Electric Combustion chamber construction
GB9106085D0 (en) 1991-03-22 1991-05-08 Rolls Royce Plc Gas turbine engine combustor
FR2723177B1 (fr) 1994-07-27 1996-09-06 Snecma Chambre de combustion comportant une double paroi
DE10155420A1 (de) 2001-11-12 2003-05-22 Rolls Royce Deutschland Hitzeschildanordnung mit Dichtungselement
DE10158548A1 (de) * 2001-11-29 2003-06-12 Rolls Royce Deutschland Brennkammerschindel für eine Gasturbine mit mehreren Kühllöchern mit unterschiedlicher Winkelausrichtung
DE10214570A1 (de) 2002-04-02 2004-01-15 Rolls-Royce Deutschland Ltd & Co Kg Mischluftloch in Gasturbinenbrennkammer mit Brennkammerschindeln
US7140185B2 (en) 2004-07-12 2006-11-28 United Technologies Corporation Heatshielded article
US9587832B2 (en) * 2008-10-01 2017-03-07 United Technologies Corporation Structures with adaptive cooling
US8359865B2 (en) 2010-02-04 2013-01-29 United Technologies Corporation Combustor liner segment seal member
US8997495B2 (en) * 2011-06-24 2015-04-07 United Technologies Corporation Strain tolerant combustor panel for gas turbine engine
US20130000309A1 (en) * 2011-06-30 2013-01-03 United Technologies Corporation System and method for adaptive impingement cooling
DE102012025375A1 (de) * 2012-12-27 2014-07-17 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Anordnung von Prallkühllöchern und Effusionslöchern in einer Brennkammerwand einer Gasturbine

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6701714B2 (en) * 2001-12-05 2004-03-09 United Technologies Corporation Gas turbine combustor
US7363763B2 (en) * 2003-10-23 2008-04-29 United Technologies Corporation Combustor
US7219498B2 (en) * 2004-09-10 2007-05-22 Honeywell International, Inc. Waffled impingement effusion method

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180299126A1 (en) * 2017-04-18 2018-10-18 United Technologies Corporation Combustor liner panel end rail
US20180306113A1 (en) * 2017-04-19 2018-10-25 United Technologies Corporation Combustor liner panel end rail matching heat transfer features

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Publication number Publication date
DE102016222099A1 (de) 2018-05-17
EP3321583B1 (fr) 2021-03-17
EP3321583A1 (fr) 2018-05-16

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