EP2527743B1 - Composant segmenté à base de matériau réfractaire pour une chambre de combustion annulaire, chambre de combustion annulaire pour un moteur d'aéronef, moteur d'aéronef et procédé de fabrication d'une chambre de combustion annulaire - Google Patents

Composant segmenté à base de matériau réfractaire pour une chambre de combustion annulaire, chambre de combustion annulaire pour un moteur d'aéronef, moteur d'aéronef et procédé de fabrication d'une chambre de combustion annulaire Download PDF

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Publication number
EP2527743B1
EP2527743B1 EP12169511.8A EP12169511A EP2527743B1 EP 2527743 B1 EP2527743 B1 EP 2527743B1 EP 12169511 A EP12169511 A EP 12169511A EP 2527743 B1 EP2527743 B1 EP 2527743B1
Authority
EP
European Patent Office
Prior art keywords
combustion
combustion chamber
chamber
chamber wall
annular
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.)
Not-in-force
Application number
EP12169511.8A
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German (de)
English (en)
Other versions
EP2527743A2 (fr
EP2527743A3 (fr
Inventor
Karl Schreiber
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
Publication of EP2527743A2 publication Critical patent/EP2527743A2/fr
Publication of EP2527743A3 publication Critical patent/EP2527743A3/fr
Application granted granted Critical
Publication of EP2527743B1 publication Critical patent/EP2527743B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F23R3/42—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
    • F23R2900/00018—Manufacturing combustion chamber liners or subparts
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49229—Prime mover or fluid pump making
    • Y10T29/4927—Cylinder, cylinder head or engine valve sleeve making

Definitions

  • the invention relates to a segment component of high-temperature casting material for an annular combustion chamber, an annular combustion chamber for an aircraft engine, an aircraft engine and a method for producing an annular combustion chamber.
  • annular combustion chambers which are arranged axially between the compressor and the turbine.
  • An annular combustion chamber has coaxially with the engine longitudinal axis a combustion chamber walls bounded annular space, which is also referred to as a flame tube.
  • the injectors for the fuel are arranged along the annular cross section of the annular space. In operation, the fuel flames extend from these injectors into the annulus.
  • a combustion chamber wall which during operation shields a fuel flame extending along a burner axis from the environment, has a bulge, the bulge pointing in a direction pointing away from the burner axis.
  • a part of a segmental component for an outer combustion chamber wall of an annular combustion chamber has e.g. a bulge that points radially outward.
  • a portion of a segmental component for an internal combustion chamber wall includes e.g. a bulge that points outward.
  • the segmental component has a combustion chamber head, an inner combustion chamber wall and an outer combustion chamber wall, between which a fuel flame is arranged along a burner axis during operation.
  • the inner combustion chamber wall, the combustion chamber head and the outer combustion chamber wall are connected together as a one-piece, U-shaped molded casting.
  • the inner and / or the outer combustion chamber wall have a bulge in the direction pointing away from the burner axis.
  • the at least one bulge of the combustion chamber wall is adapted substantially to the contour of the fuel flame during operation.
  • the length and / or width of the derecognition essentially correspond to the length and / or width of the fuel flame during operation.
  • Advantageous high temperature cast materials are a superalloy containing nickel, chromium, cobalt and / or nickel-iron, especially Inconel 738 / Inconel 738 LC, Inconel 939 / Inconel 939 LC, Inconel 713 / Inconel 713 LC, C1023, Mar M 002 and / or CM 274LC , These materials have a sufficient temperature resistance.
  • an advantageous embodiment is provided if at least one mounting flange is arranged on the combustion chamber head. Furthermore, it is advantageous if a device for arranging an injector for fuel is provided on the combustion chamber head. Also, advantageously, at least one integrally formed on a combustion chamber wall nozzle for cooling air can be provided.
  • the combustion chamber wall has an average thickness of between 1 and 4 mm, in particular 1.4 to 3 mm.
  • annular combustion chamber for an aircraft engine with the features of claim 8.
  • at least two segment components according to at least one of claims 1 to 7 are used.
  • annular combustion chamber have a variable annular space height along the circumference of the annular space.
  • the annulus height By adapting the annulus height to, for example, burner flames and / or injectors, the thermal and / or mechanical loading of the walls can be achieved. This applies in particular when regions A of a larger annular space height H RA alternate with regions B of a smaller annular space height H RB along the circumference, so that the combustion chamber walls form a type of wave-like structure
  • areas are formed with a larger annulus height and areas with a smaller annulus height, which are arranged in the assembly injectors for the fuel in the areas with the larger annulus height.
  • the areas of larger annulus height give the fuel flame more space and shield it from annoyance in the annulus.
  • the segment components are interconnected by welds, in particular by electron beam welding, laser welds with IN626 Filler, Polymet 972 or other ductile welding consumables.
  • the object is also achieved by methods for producing an annular combustion chamber.
  • At least two segment components made of high-temperature cast material having an inner combustion chamber wall, an outer combustion chamber wall and a combustion chamber head are cast.
  • the inner combustion chamber wall, the outer combustion chamber wall, and the combustion chamber head are connected together as a one-piece, U-molded casting, wherein the inner and / or outer combustion chamber wall has a bulge in the direction away from a burner axis.
  • the at least two segment components are connected by welding to the annular combustion chamber.
  • FIG. 1 In a perspective view, an annular combustion chamber is shown with an annular space 30, as used for example in an aircraft engine.
  • the annular space 30 is arranged in the main flow direction of the aircraft engine behind the compressor, not shown here, and the inlet region of a turbine 40.
  • two injectors 25 are visible, from which fuel flames 20 (not shown here) emerge along burner axes 21 during operation.
  • the burner axes 21 and thus also the fuel flames 20 are thus between the inner combustion chamber wall 11 and the outer combustion chamber wall 12.
  • This annular space 30 is also referred to as a flame tube.
  • the combustion chamber walls 11, 12 thus shield the fuel flames 20 inwardly and outwardly from the environment.
  • the annulus height H R (also referred to as Flammraum hope) varies in the axial direction of the aircraft engine, but is constant along the circumference of the annular combustion chamber 10.
  • the invention described below with reference to various embodiments relates, inter alia, annular combustion chambers in which the ring combustion chamber height H R is non-constant along the circumference.
  • Such an annular combustion chamber is e.g. composed of at least two segment components 10 made of high temperature casting material.
  • each of the segmental components 10 would be e.g. 180 ° of the annulus 30 provide.
  • a segment component 10 is shown, which covers a much smaller angular range, namely 30 °, as in the view of Fig. 2A is particularly clear.
  • An annular combustion chamber composed of such segmental components 10, therefore has 12 of these segmental components 10.
  • a segmental component 10 is shown in which parts form the inner combustion chamber wall 11 and the outer combustion chamber wall 12 when the segmental components 10 are assembled (see FIG Fig. 5 ).
  • an opening 24 is provided for the injector 25, not shown here.
  • the fuel flame 20 (not shown here) produced by the injector 25 extends along the burner axis 21 into the annular space 30 in the direction of the inlet region of the turbine 40 (not shown here, see FIG Fig. 1 ).
  • This embodiment of the segment component 10 is produced in one piece from a high-temperature casting material.
  • a superalloy may be used which contains nickel, chromium, cobalt and / or nickel-iron.
  • Typical high temperature cast alloys are in particular Inconel 738 / Inconel 738 LC, Inconel 939 / Inconel 939 LC, Inconel 713 / Inconel 713 LC, C1023, Mar M 002 and / or CM 274LC.
  • the casting processes e.g., investment casting) allow segmental components 10 to be made with very thin walls and in very complex shapes.
  • the combustion chamber walls 11, 12 have an average thickness of between 1 and 4 mm.
  • the wall of the combustion chamber head 23 can between 2 and 4 mm.
  • the shaping it is possible, for example, to form nozzle 15 for air cooling during casting.
  • mounting flanges 23 can be molded integrally on the combustion chamber head 22. Basically, the possibilities of shaping are not limited to the illustrated features.
  • the combustion chamber walls 11, 12 of this embodiment are contoured in a particular way.
  • the inner combustion chamber wall 11 has a bulge 13, which points downwards in the representation selected here.
  • the bulge 13 thus points away from the burner axis 21.
  • the outer combustion chamber wall 12 has a bulge 14, which has the same shape in a slightly upward direction.
  • the bulge 14 thus also points away from the burner axis 21.
  • the bulges 13, 14 are arranged so that they correspond approximately to the contour of the fuel flame 20 when the annular combustion chamber is in operation.
  • the bulge 13 on the inner combustion chamber wall 11 and the bulge 14 on the outer combustion chamber wall 12 extend in the axial direction approximately as far as the fuel flame 20 extends into the annular space.
  • the axial extension of the bulges, 13, 14 is about 50 to 90% of the total axial extent of the annular space. Further, it is advantageous if the width B B of the bookings 13, 14 is about 30 to 60% of the width B of a segment component 10, wherein the width B of the bulge is smaller on the inside than on the outside.
  • Fig. 2C is the one to view the Fig. 2B also shown that the bulges 13, 14 are approximately adapted to the contour of the fuel flame.
  • a region A is drawn in, in which the annulus height H RA is increased by the stakes 13, 14 and a region B, in which the annulus height H RB is reduced.
  • An arc length U of the segment component 10 is thus composed of A + 2B. It is advantageous if the proportion of the range A is 50 to 80% of the arc length U and the portion of the range B is 20 to 50% of the arc length U.
  • H RA 0.7 - 0.9 H conv . This means that the height of the combustion chamber in the area outside the bulges 13, 14 is 70 to 90% of the usual height.
  • segment components 10 are connected to each other, an annular combustion chamber is formed whose annular space height H R is variable in the circumferential direction.
  • Segment components 10 are connected to each other, for example by laser or electron beam welding, whereby the introduced path energy is minimized. It can be a suitable, ductile filler used in welding (IN625 or Polymet 972).
  • Such a compound rivet combustion chamber is in Fig. 3 shown.
  • segment components 10 are used here to form an annular space 30.
  • Areas A of a larger annular space height H RA alternate with areas B of a smaller annular space height H RB along the circumference, so that the combustion chamber walls 11, 12 form a kind of wave-like structure.
  • the fuel flames 20 (not shown here) are in each case in the extended areas A. Between the fuel flames 20 are narrowed areas B. This leads to the effect that each fuel flame 20 can effectively burn in its own combustion chamber. Disturbances in a region of the annular space 30 can spread more severely in the entire annular space 30 due to the constrictions in the regions B.
  • air may be directed from the compressor to the turbine 40 with less severe deflection, thereby reducing the pressure loss on this flow path.
  • the described embodiment also has advantageous effects outside of the annular space 30, since the turbine cooling air K, which is guided outside the annular space, is also influenced by the contouring of the combustion chamber walls 11, 12.
  • the pressure loss during the transfer of the turbine cooling air K from the compressor outlet to the combustion chamber is determined to enter the cooling system by the flow guidance in this way. If the turbine cooling air K has to be deflected repeatedly (in particular radially) and accelerated (and then decelerated again), then the pressure loss increases. In the burner axis 21, only a small amount of turbine cooling air K flows past the burner and mixing air hole in the direction of the turbine, so the pressure loss is not so decisive there.
  • the combustion chamber head 22 is designed so that the turbine cooling air K is not greatly deflected radially outwards and inwards. These are the areas B between the bulges 13, 14, but at the respective outer sides of the annular space 30. After the radial deflection then takes place a deflection in the axial direction. Thus, in area B, there is a small deflection into the much deeper annuli around the narrower combustion chamber at this point.
  • the flow of turbine cooling air K is in Fig. 3 shown schematically.
  • the bulges 13, 14 cause a more uniform temperature distribution in the circumferential direction to form in the combustion chamber walls 11, 12, which has a positive influence on the service life of the annular combustion chamber.
  • the combustion chamber wall 11, 12, due to the bulges 13, 14 relatively far away from the fuel flame 20.
  • the combustion chamber walls 11, 12 are closer together, since the annulus height H R is lower here.
  • the wall regions of the combustion chamber walls 11, 12 which are closest to the fuel flame 20 would be hotter than other regions. For these reasons, not so much cooling air needs to be used in area A. The cooling air thus saved is available for measures to reduce exhaust emissions.
  • the inner combustion chamber wall 11 and the outer combustion chamber wall 12 have a wavy structure when made of segmental components 10, for example according to FIG Fig. 2 are composed.
  • This wavy structure allows for easier compensation of thermal and / or mechanical stresses in the combustion chamber walls 11, 12 than would be possible in annular spaces with circular cross-sections in the circumferential direction.
  • the segmental components 10 may be provided with a thermal barrier coating.
  • a further embodiment of a segment component 10 is shown. Basically, it has the same functions and properties as the segment component 10 described above, so that reference can be made to the corresponding description.
  • the bulge 13 has a rather small width in the vicinity of the combustion chamber head 23, which widens steadily, in order then to become smaller again.
  • a segment component 10 it is also possible for a segment component 10 to have only one outer or inner part of the annular combustion chamber.
  • Fig. 5 an embodiment of a segment component 10 is shown, which has only one outer combustion chamber wall 12. Like the previously described embodiments, this segmental component 10 also has a bulge 14 that faces away from the burner axis 21. To illustrate the use of this segmental component 10, are in Fig. 5 dashed lines the fuel flame 20 and the burner axis 21 drawn.
  • annular combustion chamber can be constructed, as shown in FIG Fig. 6A , B is shown.
  • segment components 10 are connected to an inner full ring structure 31, in particular welded.
  • segment components 10 are connected to an outer full ring structure 32, in particular welded Fig. 6A are the two full ring structures 31, 32 shown, each having only six segment components 10 for reasons of simplicity.
  • the inner full ring structure 31 and the outer full ring structure 32 are connected to a combustion head structure 43, as shown in FIG Fig. 6B is shown.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Laser Beam Processing (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)

Claims (13)

  1. Composant de segment en matériau moulé à haute résistance thermique pour une chambre de combustion annulaire d'un réacteur d'avion, lequel composant présente une paroi de chambre de combustion (11, 12) qui en fonctionnement isole du milieu ambiant une flamme de combustible (20) s'étendant le long d'un axe de brûleur (21),
    - sachant que la paroi de chambre de combustion (11, 12) présente un renflement (13, 14) dans un sens opposé à l'axe de brûleur (21),
    - que le composant de segment présente une tête de chambre de combustion (22), une paroi intérieure de chambre de combustion (11) et une paroi extérieure de chambre de combustion (12) entre lesquelles en fonctionnement est placée la flamme de combustible (20), et
    caractérisé en ce que la paroi intérieure de chambre de combustion (11), la tête de chambre de combustion (22) et la paroi extérieure de chambre de combustion (12) sont reliées entre elles en tant que pièce moulée monobloc en forme de U.
  2. Composant de segment selon la revendication n° 1, caractérisé en ce que la paroi intérieure et/ ou la paroi extérieure de chambre de combustion (11, 12) présente(nt) un renflement (13, 14) dans le sens opposé à l'axe de brûleur (21).
  3. Composant de segment selon la revendication n° 1 ou n° 2, caractérisé en ce que l'au moins un renflement (13, 14) de la paroi de chambre de combustion (11, 12) est pour l'essentiel adapté au contour de la flamme de combustible (20) en fonctionnement, en particulier que la paroi de chambre de combustion (11, 12) présente un renflement (13, 14) dont la longueur (LB) et/ ou la largeur (BB) correspond(ent) sensiblement à la longueur et/ ou à la largeur de la flamme de combustible (20) en fonctionnement.
  4. Composant de segment selon au moins une des revendications précédentes, caractérisé en ce que le matériau moulé à haute résistance thermique est un superalliage contenant du nickel, du chrome, du cobalt et/ ou du nickel-fer, en particulier Inconel 738/ Inconel 738 LC, Inconel 939/ Inconel 939 LC, Inconel 713/ Inconel 713 LC, C1023, Mar M 002 et/ ou CM 274LC.
  5. Composant de segment selon au moins une des revendications précédentes, caractérisé en ce que sur la tête de chambre de combustion (22) est disposée au moins une bride de fixation (23) et/ ou que sur la tête de chambre de combustion (22) est prévu un dispositif (24) pour placer un injecteur (25).
  6. Composant de segment selon au moins une des revendications précédentes, caractérisé par au moins une tubulure (15) pour l'air de refroidissement formée intégralement dans une paroi de chambre de combustion (11, 12).
  7. Composant de segment selon au moins une des revendications précédentes, caractérisé en ce que la paroi de chambre de combustion (11, 12) présente une épaisseur moyenne comprise entre 1 et 4 mm, en particulier entre 1,4 et 3 mm.
  8. Chambre de combustion annulaire pour un réacteur d'avion doté d'au moins deux composants de segment (10) selon au moins une des revendications n° 1 à n° 7.
  9. Chambre de combustion annulaire selon la revendication n° 8, caractérisée par une hauteur d'espace annulaire (HR) variable le long de la circonférence de l'espace annulaire (30), sachant qu'en particulier des zones (A) d'une plus grande hauteur d'espace annulaire (HRA alternent avec des zones (B) d'une plus faible hauteur d'espace annulaire (HRB) le long de la circonférence.
  10. Chambre de combustion annulaire selon la revendication n° 8 ou n° 9, caractérisée par des zones d'une plus grande hauteur d'espace annulaire (HRA) et des zones d'une plus faible hauteur d'espace annulaire (HRB), sachant que dans l'assemblage, des injecteurs (25) pour le combustible sont placés dans les zones ayant la plus grande hauteur d'espace annulaire (HRA).
  11. Chambre de combustion annulaire selon au moins une des revendications n° 8 à n° 10, caractérisée en ce que les composants de segment (10) sont reliés entre eux par des cordons de soudure, en particulier par soudage par bombardement électronique, par des cordons de soudage au laser avec métaux d'apport IN626, Polymet 972 ou autres métaux d'apport ductiles.
  12. Réacteur d'avion doté d'une chambre de combustion annulaire selon les revendications n° 8 à n° 11.
  13. Procédé de fabrication d'une chambre de combustion annulaire selon au moins une des revendications n° 8 à n° 11, ledit procédé comprenant les étapes suivantes:
    a) moulage d'au moins deux composants de segment (10) en matériau moulé à haute résistance thermique, avec une paroi intérieure de chambre de combustion (11), une paroi extérieure de chambre de combustion (12) et une tête de chambre de combustion (22), lesquelles sont reliées entre elles en tant que pièce moulée monobloc en forme de U, sachant que la paroi intérieure et/ ou la paroi extérieure de chambre de combustion (11, 12) présente(nt) un renflement (13, 14) dans le sens opposé à un axe de brûleur (21), et ensuite
    b) soudage des au moins deux composants de segment (10) pour former une chambre de combustion annulaire (10).
EP12169511.8A 2011-05-25 2012-05-25 Composant segmenté à base de matériau réfractaire pour une chambre de combustion annulaire, chambre de combustion annulaire pour un moteur d'aéronef, moteur d'aéronef et procédé de fabrication d'une chambre de combustion annulaire Not-in-force EP2527743B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102011076473A DE102011076473A1 (de) 2011-05-25 2011-05-25 Segmentbauteil aus Hochtemperaturgussmaterial für eine Ringbrennkammer, Ringbrennkammer für ein Flugzeugtriebwerk, Flugzeugtriebwerk und Verfahren zur Herstellung einer Ringbrennkammer

Publications (3)

Publication Number Publication Date
EP2527743A2 EP2527743A2 (fr) 2012-11-28
EP2527743A3 EP2527743A3 (fr) 2015-01-21
EP2527743B1 true EP2527743B1 (fr) 2016-09-28

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US (1) US8646279B2 (fr)
EP (1) EP2527743B1 (fr)
DE (1) DE102011076473A1 (fr)

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DE102013222863A1 (de) 2013-11-11 2015-05-13 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer sowie Verfahren zu deren Herstellung
DE102014204468A1 (de) 2014-03-11 2015-10-01 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer sowie Verfahren zu deren Herstellung
DE102016201452A1 (de) * 2016-02-01 2017-08-03 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit Wandkonturierung
CN111059575B (zh) * 2018-10-16 2022-05-10 中发天信(北京)航空发动机科技股份有限公司 涡喷发动机火焰筒外壳
US20200318549A1 (en) * 2019-04-04 2020-10-08 United Technologies Corporation Non-axisymmetric combustor for improved durability
GB202019219D0 (en) * 2020-12-07 2021-01-20 Rolls Royce Plc Lean burn combustor
GB202019222D0 (en) 2020-12-07 2021-01-20 Rolls Royce Plc Lean burn combustor
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EP2100687A1 (fr) * 2008-02-29 2009-09-16 Siemens Aktiengesellschaft Chauffage de fil en acier sans potentiel lors de la soudure et son dispositif

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EP2527743A2 (fr) 2012-11-28
US20120304658A1 (en) 2012-12-06
US8646279B2 (en) 2014-02-11
EP2527743A3 (fr) 2015-01-21
DE102011076473A1 (de) 2012-11-29

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