EP3034944A1 - Chambre de combustion de turbines a gaz a epaisseur de paroi variable - Google Patents

Chambre de combustion de turbines a gaz a epaisseur de paroi variable Download PDF

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Publication number
EP3034944A1
EP3034944A1 EP15198122.2A EP15198122A EP3034944A1 EP 3034944 A1 EP3034944 A1 EP 3034944A1 EP 15198122 A EP15198122 A EP 15198122A EP 3034944 A1 EP3034944 A1 EP 3034944A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
gas turbine
air holes
turbine combustor
wall
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.)
Withdrawn
Application number
EP15198122.2A
Other languages
German (de)
English (en)
Inventor
Dr.-Ing. Miklós Gerendás
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
Publication of EP3034944A1 publication Critical patent/EP3034944A1/fr
Withdrawn legal-status Critical Current

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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
    • F23R3/02—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
    • 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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • 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
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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
    • F23R2900/00005—Preventing fatigue failures or reducing mechanical stress in gas turbine components

Definitions

  • the invention relates to a gas turbine combustor according to the preamble of claim 1.
  • the invention relates to a gas turbine combustor having at least one combustion chamber wall in which mixing air holes are formed in predetermined regions.
  • the invention has for its object to provide a gas turbine combustor of the type mentioned above, which avoids the disadvantages of the prior art with a simple structure and simple, cost manufacturability and in particular in the area of the mixing air holes has sufficient strength.
  • the combustion chamber has a greater wall thickness in the area of the mixed air holes than in the areas not provided with mixed air holes.
  • the mixing air holes are arranged in a central region of the combustion chamber, based on the axial extent of the combustion chamber and in the circumferential direction of the combustion chamber.
  • This annular peripheral region, in which the mixing air holes are arranged, is provided according to the invention with a greater wall thickness.
  • the load-bearing wall thickness of the combustion chamber wall is thus increased locally to the extent in which the cross section carrying the mixing air holes in the circumferential direction is removed. According to the invention, it is possible to use this solution both in single-layer and in two-layer combustion chamber walls. In the case of a two-layer combustion chamber wall, it is possible according to the invention to provide or to thicken only one layer, for example the load-bearing outer combustion-chamber wall or the hot, inner combustion-chamber wall or both.
  • the stiffness of the combustion chamber wall no longer varies in the longitudinal direction, but is constant, in particular in the area of the mixing air holes, in particular compared with the areas in which no mixed air holes are formed.
  • the external load distortions do not concentrate in the area provided with the mixed air holes.
  • a gap which may arise between the shingle (inner combustion chamber wall) and the shingle support (outer combustion chamber wall), becomes smaller due to the constant rigidity.
  • the invention is applicable both to combustion chamber walls which are manufactured as a cast part, and to combustion chamber walls which are produced by means of a generative method (laser sintering, ALM, additive layer manufacturing).
  • ALM laser sintering
  • ALM additive layer manufacturing
  • the invention can also apply to combustor walls made of fiber reinforced ceramic (CMC).
  • CMC fiber reinforced ceramic
  • the number of layers of ceramic fiber fabric or windings in the area of the mixed air holes is increased locally. This is associated with little additional effort, since the wall is basically constructed of several layers, only in the area of the mixed air holes, the number of layers of, for example, 12 increased to 20.
  • the increase in wall thickness is more pronounced in CMC, since the wall temperature may be higher, so less cooling air must be used, and thus more air to be passed through the mixing holes, which increases their diameter beyond what is possible with a metallic construction.
  • the layers can be inserted on the inside or outside or as additional intermediate layers with limited axial extent.
  • the shingle can be thickened by eliminating the bridge to the outside. This can also be done in the form of a rib on the cold side of the shingle, which then protrudes through the opening in the cold combustion chamber wall.
  • the gas turbine engine 110 is a generalized example of a turbomachine, in which the invention can be applied.
  • the engine 110 is formed in a conventional manner and comprises in succession an air inlet 111, a fan 112 circulating in a housing, a medium pressure compressor 113, a high pressure compressor 114, a combustion chamber 115, a high pressure turbine 116, a medium pressure turbine 117 and a low pressure turbine 118 and a Exhaust nozzle 119, which are all arranged around a central engine center axis 101.
  • the intermediate pressure compressor 113 and the high pressure compressor 114 each include a plurality of stages, each of which includes a circumferentially extending array of fixed stationary vanes 120, commonly referred to as stator vanes, which are radially inwardly of the engine casing 121 in an annular manner Flow channel through the compressors 113, 114 protrude.
  • the compressors further include an array of compressor blades 122 projecting radially outward from a rotatable drum or disk 125 coupled to hubs 126 of high pressure turbine 116 and intermediate pressure turbine 117, respectively.
  • the turbine sections 116, 117, 118 have similar stages, comprising an array of fixed vanes 123 projecting radially inward from the housing 121 into the annular flow passage through the turbines 116, 117, 118, and a downstream array of turbine blades 124, projecting outwardly from a rotatable hub 126.
  • the compressor drum or compressor disk 125 and the vanes 122 disposed thereon and the turbine rotor hub 126 and turbine blades 124 disposed thereon rotate about the engine centerline 101 during operation.
  • the Fig. 2 shows a longitudinal sectional view of a known from the prior art combustion chamber wall in an enlarged view.
  • a combustion chamber 1 is shown with a central axis 9, which comprises a combustion chamber head 3, a base plate 8 and a heat shield 2.
  • a burner seal is provided with the reference numeral 4.
  • the combustion chamber 1 has an outer cold combustion chamber wall 7, to which an inner, hot combustion chamber wall 6 is attached.
  • mixed air mixing air holes 5 are provided for supplying mixed air mixing air holes 5 are provided. The presentation of impingement cooling holes and effusion holes has been omitted for clarity.
  • the inner combustion chamber wall 6 is provided with bolts 13, which are designed as threaded bolts and are screwed by means of nuts 14.
  • the storage of the combustion chamber 1 via Brennschflansche 12 and 11 Brennschhfitmaschine.
  • Prior art combustor walls made of sheet metal typically have a constant thickness in the range of 0.9 to 1.6 mm, while combustor walls made as castings have wall thicknesses of between 1.2 and 2.5 mm have.
  • the 3 and 4 show in a schematic side view of a combustion chamber wall, the arrangement of mixed air holes.
  • the Fig. 3 shows, for example, the assignment of mixed air holes, as they are known from the prior art. It will be understood that the change in wall thickness, and hence rigidity, of the combustion chamber wall will depend on the arrangement and pattern of the mixing air holes. In particular, the axial distance and the circumferential distance of the mixed air holes must be taken into account. Furthermore, the respective diameters of the mixed air holes play a role.
  • the wall thickness is increased substantially by 41%.
  • the wall thickness is increased substantially by 73%.
  • the smallest web width occurs between the mixed air holes of one row or between the mixed air holes of adjacent rows. This only determines the axial position of the maximum wall thickness. If the smallest web width lies between the mixed air holes of a row, then the maximum of the wall thickness lies at the axial position of the axes of the mixed air hole row. If the minimum land width lies between mixed air holes of adjacent rows of mixed air holes, then the maximum wall thickness between the central axes of the two rows of mixing air holes is substantially midway between the rows.
  • the axial extent of the thickening for a mixed air hole row is substantially limited to the range between a hole diameter upstream and a hole diameter downstream.
  • the axial extent of the thickening for two rows of mixed air holes is limited to the range between a hole diameter of the upstream mixed air hole row upstream and a hole diameter of the downstream mixed air hole row downstream.
  • the wall thickness may be increased in a ramp in front of the thickness-determining ligament, followed by a region of constantly high wall thickness in the area of the mixing air holes and a ramp back to a lesser wall thickness, which then substantially until shortly before the end of the combustion chamber is maintained.
  • the substantially constant wall thickness before the mixing air hole row need not be identical to the substantially constant wall thickness downstream thereof.
  • the transitions in the wall thickness are fluent designed to avoid voltage spikes by cross-sectional jumps.
  • the sheet thickness of the outer, cold combustion chamber wall 7 from 1.2 mm to 1.6 mm, while the thickness of a formed as a casting, inner, hot combustion chamber wall 6 in the mixing air holes of 1.4 mm 2 mm is increased.
  • the Fig. 4 illustrates a possible embodiment according to the invention, in which the two Mischluftlochschloisn are substantially approximated in the circumferential direction or almost overlap.
  • the weakening of the combustion chamber wall would be further increased. According to the invention thus takes place in this area a greater thickening, as it follows in connection with the Fig. 5 and 6 is described.
  • a sheet thickness of a combustion chamber wall can be increased from 1.2 mm to 1.8 mm.
  • the wall thickness of a casting of 1.4 mm can be increased according to the invention, for example in the region of the intersection of the mixed air holes to 2.5 mm.
  • the invention can be used both with single-walled and two-walled combustion chambers.
  • a single-walled combustion chamber for example, the wall thickness of the metal sheet in the area of the mixed air holes is increased or the adjacent area not provided with mixed air holes is reduced in its cross-section by ironing pressures.
  • Abstreckrise a move away from standardized sheet thicknesses to a adapted to the local requirements wall thickness.
  • By means of ironing produced components with locally adapted wall thickness are cheaper to produce than components which are joined from several sheets, forgings or castings.
  • a multilayer wall construction of the combustion chamber wall which is produced for example as a cassette or by joining laminated sheets, it is possible according to the invention to adapt the wall thickness in the region of the mixing air holes analogously to the local requirements.
  • the Fig. 5 shows a sectional view of a combustion chamber analogous to Fig. 2
  • Fig. 2 is in the upper part of the figure as a diagram the wall thickness W over the length of the Combustion chamber wall applied. It follows that the combustion chamber wall has a constant wall thickness W over its entire length.
  • the in Fig. 5 shown embodiment in the area of the mixing air holes 5, a thickening of the wall and thus a greater wall thickness provided, as can be seen from the diagram in the upper half of the figure Fig. 5 results.
  • the Fig. 5 shows in the sectional view of a construction in which both the outer, cold combustion chamber wall 7, and the inner, hot combustion chamber wall 8 are formed thickened.
  • the Fig. 6 shows a further embodiment in a representation analog Fig. 5 , It can be seen that in addition to the in Fig. 5 provided thickening or increasing the wall thickness in the area between the mixing air holes 5 and in the region of their overlap (s Fig. 4 ) a further increase in wall thickness or wall thickness occurs. This results in particular from the diagram in the upper half of the representation of Fig. 6 ,

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP15198122.2A 2014-12-19 2015-12-04 Chambre de combustion de turbines a gaz a epaisseur de paroi variable Withdrawn EP3034944A1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102014226707.2A DE102014226707A1 (de) 2014-12-19 2014-12-19 Gasturbinenbrennkammer mit veränderter Wandstärke

Publications (1)

Publication Number Publication Date
EP3034944A1 true EP3034944A1 (fr) 2016-06-22

Family

ID=54783486

Family Applications (1)

Application Number Title Priority Date Filing Date
EP15198122.2A Withdrawn EP3034944A1 (fr) 2014-12-19 2015-12-04 Chambre de combustion de turbines a gaz a epaisseur de paroi variable

Country Status (3)

Country Link
US (1) US20160178198A1 (fr)
EP (1) EP3034944A1 (fr)
DE (1) DE102014226707A1 (fr)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10267521B2 (en) * 2015-04-13 2019-04-23 Pratt & Whitney Canada Corp. Combustor heat shield
DE102016219424A1 (de) 2016-10-06 2018-04-12 Rolls-Royce Deutschland Ltd & Co Kg Brennkammeranordnung einer Gasturbine sowie Fluggasturbine
DE102021214499A1 (de) 2021-12-16 2023-06-22 Rolls-Royce Deutschland Ltd & Co Kg Brennkammerbaugruppe mit spezifisch angeordneten Mischluftlöchern an innerer und äußerer Brennkammerwand

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20020116929A1 (en) 2001-02-26 2002-08-29 Snyder Timothy S. Low emissions combustor for a gas turbine engine
WO2003042597A1 (fr) * 2001-11-15 2003-05-22 Siemens Aktiengesellschaft Chambre de combustion annulaire pour turbine a gaz
US20030131603A1 (en) * 2002-01-16 2003-07-17 Bolender Lynn Marie Method and apparatus for relieving stress in a combustion case in a gas turbine engine
US20040074239A1 (en) * 2002-10-21 2004-04-22 Peter Tiemann Annular combustion chambers for a gas turbine and gas turbine
EP1528322A2 (fr) 2003-10-23 2005-05-04 United Technologies Corporation Chambre de combustion
EP1795809A2 (fr) 2005-12-06 2007-06-13 United Technologies Corporation Chambre de combustion de turbine à gaz
US20120240584A1 (en) * 2009-12-11 2012-09-27 Snecma Combustion chamber for a turbine engine
WO2014160299A1 (fr) * 2013-03-14 2014-10-02 United Technologies Corporation Panneau de chambre de combustion à durée de vie accrue

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4738902A (en) * 1983-01-18 1988-04-19 United Technologies Corporation Gas turbine engine and composite parts
US5233828A (en) * 1990-11-15 1993-08-10 General Electric Company Combustor liner with circumferentially angled film cooling holes
GB9803291D0 (en) * 1998-02-18 1998-04-08 Chapman H C Combustion apparatus
GB2373319B (en) * 2001-03-12 2005-03-30 Rolls Royce Plc Combustion apparatus
US6675582B2 (en) * 2001-05-23 2004-01-13 General Electric Company Slot cooled combustor line
US6895757B2 (en) * 2003-02-10 2005-05-24 General Electric Company Sealing assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor
US8661826B2 (en) * 2008-07-17 2014-03-04 Rolls-Royce Plc Combustion apparatus
US9062884B2 (en) * 2011-05-26 2015-06-23 Honeywell International Inc. Combustors with quench inserts

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20020116929A1 (en) 2001-02-26 2002-08-29 Snyder Timothy S. Low emissions combustor for a gas turbine engine
WO2003042597A1 (fr) * 2001-11-15 2003-05-22 Siemens Aktiengesellschaft Chambre de combustion annulaire pour turbine a gaz
US20030131603A1 (en) * 2002-01-16 2003-07-17 Bolender Lynn Marie Method and apparatus for relieving stress in a combustion case in a gas turbine engine
US20040074239A1 (en) * 2002-10-21 2004-04-22 Peter Tiemann Annular combustion chambers for a gas turbine and gas turbine
EP1528322A2 (fr) 2003-10-23 2005-05-04 United Technologies Corporation Chambre de combustion
EP1795809A2 (fr) 2005-12-06 2007-06-13 United Technologies Corporation Chambre de combustion de turbine à gaz
US20120240584A1 (en) * 2009-12-11 2012-09-27 Snecma Combustion chamber for a turbine engine
WO2014160299A1 (fr) * 2013-03-14 2014-10-02 United Technologies Corporation Panneau de chambre de combustion à durée de vie accrue

Also Published As

Publication number Publication date
DE102014226707A1 (de) 2016-06-23
US20160178198A1 (en) 2016-06-23

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