• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    COMBUSTION CHAMBER FOR TURBINE ENGINE. The invention relates to a combustion chamber (1) for a turbine such as a turbo-reactor or aircraft turboprop, comprising internal and external annular walls (3,4) of revolution, connected by an annular wall at the bottom of the chamber ( 5). The inner wall (3) consists of a single layer of material whose thickness and / or nature varies along the longitudinal axis and / or the circumferential direction of said wall (3).
    公开号:BR112012014057B1
    申请号:R112012014057-4
    申请日:2010-12-02
    公开日:2020-12-08
    发明作者:Caroline Jacqueline Denise Berdou；Laurent Bernard Cameriano
    申请人:Snecma；
    IPC主号:
    专利说明:

    [0001] The invention relates to a combustion chamber for a turbomachinery, such as turbo-reactor or aircraft turboprop.
    [0002] A combustion chamber comprises coaxial walls of revolution which extend inside each other and which are connected at their ends through an annular back wall of the chamber comprising air supply openings and means of conducting fuel with special support injectors.
    [0003] The inner and outer walls of the chamber contain holes for the entry of primary air and dilution air, and areas with multi-perforations for the passage of cooling air.
    [0004] To better withstand extreme temperatures, it is known to mount thermal barriers on the walls of the combustion chamber, these barriers appearing in the form of additional material thicknesses supported on the walls.
    [0005] JP 6167245 describes a combustion chamber, whose internal wall has a constant thickness and is covered with a thermal barrier of varying thickness.
    [0006] The use of a thermal barrier increases the resistance of the chamber to high temperatures, but increases its weight.
    [0007] In order to satisfy market requirements, it is necessary to reduce the weight of the combustion chamber. However, the service life of the combustion chamber must not be shortened. In particular, walls must be dimensioned to withstand creep damage. Remember that creep is the irreversible deformation of a material subjected to constant stress for a sufficient time. This deformation is accentuated by the high temperatures to which the combustion chamber walls are subjected.
    [0008] The purpose of the invention is especially to provide the solution of a problem in a simple, efficient and economical way.
    [0009] For this purpose, a combustion chamber for turbomachinery is proposed, such as a turbo-reactor or airplane turboprop, comprising internal and external annular walls of revolution, connected by an annular chamber bottom wall, characterized by the fact of its the inner wall consists of a single thickness of material, the thickness and / or nature of which varies along the longitudinal axis of the circumferential direction of said wall, while its outer annular wall has a substantially constant thickness.
    [0010] The invention allows to increase the resistance of the combustion chamber at extreme temperatures, without the use of thermal barriers, and without increasing the mass, locally modifying the thickness and / or the material nature of the chamber walls.
    [0011] The external annular wall is generally less hot than the internal annular wall and, therefore, does not require any special adaptation of its structure.
    [0012] According to an embodiment of the invention, the internal wall of the combustion chamber, consisting of a single thickness of material, has at least one zone called hot with a high thermal gradient, of greater thickness and at least one zone said cold with lower thermal gradient, less thickness.
    [0013] "Hot" zones, being zones subject to the greatest thermal gradients, it is advantageous to increase their thickness.
    [0014] According to another characteristic of the invention, the internal wall of the combustion chamber, consisting of a single thickness of material, has at least two adjacent zones or zones made up of different materials.
    [0015] As mentioned above, it is also possible to use a more resistant material locally, in the hottest areas or in the areas subjected to the highest thermal gradients, and a less resistant and lighter material, in the coldest areas or areas subjected to the thermal gradients lower.
    [0016] Preferably, the internal wall of the combustion chamber, of variable thickness, is made by machining.
    [0017] Machining allows to obtain lower dimensional tolerances than those of plate formation conventionally used for the realization of combustion chambers.
    [0018] On the other hand, machining allows to vary the thickness of the inner wall, both along the longitudinal axis and along a circumferential direction.
    [0019] Alternatively, the internal wall of the combustion chamber, of varying thickness, is made by stretching and plate formation.
    [0020] This method is simpler and less expensive than machining.
    [0021] The zones of thickness and / or variable nature of the internal wall of the combustion chamber comprise at least one of the zones belonging to the group comprising the zones located between the injectors, the zones that comprise holes of primary air, of dilution air, the zones with annular fixation flanges, and the zones with multi-perforations.
    [0022] The invention also relates to a turbomachinery, such as a turboprop or airplane turboprop, having a combustion chamber of the type described above.
    [0023] The invention will be better understood and other details, characteristics and advantages of the invention appear in the reading of the following description, given as a non-limiting example with reference to the attached drawings, in which: - Figure 1 is a partial schematic view in axial section of an annular combustion chamber of a turbomachinery. - Figure 2 is a perspective view of a sector of the combustion chamber of Figure 1. Figure 3 is a view of a detail of the invention, showing a section of the internal annular wall of the combustion chamber of Figure 1.
    [0024] As shown in Figures 1 and 2, an annular combustion chamber for turbomachinery 1 is arranged at the outlet of a diffuser 2, which is located at the outlet of a compressor (not shown), and includes annular walls of internal and external revolution 3, 4 connected upstream to an annular wall at the bottom of the chamber 5 and fixed downstream by internal and external annular flanges 6,7 respectively on an internal taper plug 8 of the diffuser 2 and on one end of an external housing 9 of the chamber 1 , the upstream end of the sump 9 being connected by an external taper plug 10 to the diffuser 2.
    [0025] The annular wall at the bottom of the chamber 5 contains openings 11 (Figure 2) through which the air from the diffuser 2 passes and which are used for the assembly of the fuel injectors 12 fixed in the outer casing 9 and regularly distributed around the longitudinal axis of the chamber. Each injector 12 includes a fuel injection head 13 centered on the opening 11 of the annular wall 5 and connected to the axis of the opening 11.
    [0026] A part of the air flow provided by the compressor and exiting through the diffuser 2 passes through the openings 11 and feeds the combustion chamber, the other part of the air flow feeding the inner and outer annular channels 14, 15 of the contour of the combustion chamber.
    [0027] The inner channel 14 is formed between the inner plug 8 of the diffuser 2 and the inner wall 3 of the chamber, and the air that passes through that channel is divided into a flow into the chamber through air holes 16, 17 primer and dilution air (Figure 2) of the inner wall 3 and flowing through the holes in the inner flange 6 of chamber 1 to cool the components, not shown, located downstream of chamber 1.
    [0028] The external channel 15 is formed between the external sump 9 and the outer wall 4 of the chamber 1, and the air that passes through this channel is divided into a flow into the chamber through holes 18, 19 of primary air and dilution air (Figure 2) from the outer wall 4 and a flow that passes through holes 7 to the outer flange to cool downstream components.
    [0029] The primary air inlet holes 16, 18 are regularly distributed over the circumferences of the inner and outer walls 3, 4, respectively, centered on the axis of the chamber 1, and the holes 17, 19 of the dilution air inlet are regularly distributed over circumferences of the inner and outer walls 3, 4, respectively, centered on the axis of the chamber 1 downstream of the holes 16, 18.
    [0030] The internal and external annular walls 3, 4 also comprise microperforations (not visible) for the passage of cooling air.
    [0031] In operation, the external and internal annular walls 3, 4 have zones having different temperatures, this temperature heterogeneity is represented schematically in Figure 2 in the form of zones 20, 21, 22, 23, distinct from each other.
    [0032] This phenomenon concerns in particular the internal annular wall 3. The temperature zones are numbered by increasing temperature values. Thus, zones 20 are relatively "cooler" zones subjected to lower thermal gradients and zone 23 is the "warmer" zone, subjected to higher thermal gradients. This division of zones is only given by way of example and results notably the particular structure of the combustion chamber 1.
    [0033] The presence and location of the different zones 20 to 23 can be highlighted by simulation by calculating or applying a temperature-reactive paint, the paint of which, after the combustion chamber works, varies locally according to the temperature.
    [0034] According to the invention, the inner wall 3 consists of a single thickness of material whose thickness and / or nature varies along the longitudinal axis and / or the circumferential direction of said wall.
    [0035] In the embodiment shown in all the figures, the thickness of the internal wall is locally variable, likewise, comprising zones 20 to 23 of different temperatures.
    [0036] Thus, as shown in Figure 3, the internal annular wall 3 is made up of a single thickness of material and comprises the zones of greater thickness e1 (see Figure 3), for example, zones 22 and 23, and zones thinner e2, for example, zones 20 and 21.
    [0037] The thickest areas are those that, in operation, are subject to the highest temperatures, for example, in the order of 1000 ° C. These zones have a thickness e1 between 1 and 2 mm, preferably in the order of 1.5 mm. Conversely, the areas with the lowest thickness are those that in operation are subject to the lowest temperatures. These zones have a thickness e2 between 0.5 and 1 mm, preferably on the order of 1 mm.
    [0038] The external annular wall 4 has a substantially constant thickness, between 1 and 1.5 mm, preferably in the order of 1.2 mm.
    [0039] It is possible, for example, starting from a known combustion chamber whose walls of revolution have a constant thickness of 1.5 mm, to make a lighter combustion chamber, having an external wall of revolution with a thickness of 1 , 2 mm and an internal wall of revolution having a thickness of 1.5 mm in the hot zones and 1 mm in the coldest zones, the mass of this chamber being that of a chamber having walls with a constant thickness equal to 1.2 mm .
    [0040] The combustion chamber according to the invention, in particular the inner wall 3 of varying thickness, is produced by machining.
    [0041] Alternatively, the internal wall 3 of variable thickness is made by stretching and forming the plate.
    [0042] According to an embodiment not shown in the drawings, the zones of varying thickness could be replaced or could include zones of different types, in order to accommodate the zones constituted of a material with high thermal resistance in the hottest zones and zones made of a material with lower thermal resistance, but lighter in the coldest zones.
    [0043] Likewise, areas of different natures may prevent the formation of cracks, the material being able to be exchanged locally so that the areas initially used under tension, in which cracks may start, are requested in compression due to the behavior of neighboring areas.
    [0044] Each of these embodiments allows to reduce the weight of the combustion chamber, improving its thermal resistance and, thus, its useful life.
    [0045] The zones of thickness and / or variable nature of the internal wall 3 are notably the zones located between the injectors 12, the zones comprising the holes of primary air 16 and of dilution air 17, the zones comprising the annular fixing flanges 6, and the zones with multi-perforations.
    权利要求:
    Claims (7)
    [0001]
    1. Combustion chamber (1), for turbomachinery such as a turboprop or airplane turboprop, comprising internal and external annular walls (3, 4) of revolution, connected by an annular chamber bottom wall (5), characterized by the fact of its internal wall (3) being constituted by a single thickness of material, whose thickness (e1, e2) and / or the nature varies along the longitudinal axis and the circumferential direction of said wall (3) while its external annular wall (4) has a substantially constant thickness and a material of a constant nature.
    [0002]
    2. Combustion chamber (1), according to claim 1, characterized in that its internal wall consisting of a single thickness of material comprises at least one so-called hot zone (23, 22) with a higher thermal gradient, of greater thickness (e1), and at least one so-called cold zone (21, 20) with a lower thermal gradient, of less thickness (e2).
    [0003]
    3. Combustion chamber (1), according to claim 1 or 2, characterized by the fact that its internal wall consisting of a single material thickness has at least two adjacent zones made up of different materials.
    [0004]
    4. Combustion chamber (1) according to any one of claims 1 to 3, characterized in that its internal wall of variable thickness is made by machining.
    [0005]
    5. Combustion chamber (1) according to any one of claims 1 to 3, characterized in that its internal wall of variable thickness is made by drawing and forming a plate.
    [0006]
    6. Combustion chamber (1) according to any one of claims 1 to 5, characterized in that the zones of thickness (e1, e2) and / or of variable nature of its internal wall comprise at least one of the zones that they are part of the group comprising the zones located between the injectors (12), the zones that comprise holes of primary air (16, 18) and of dilution air (17, 19), the zones comprising annular fixing flanges (6, 7 ), and zones with multi-perforations.
    [0007]
    7. Turbomachinery, such as a turboprop or airplane turboprop, characterized by the fact that it comprises a combustion chamber (1) as defined in any one of claims 1 to 6.
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    FR2922630B1|2007-10-22|2015-11-13|Snecma|COMBUSTION CHAMBER WALL WITH OPTIMIZED DILUTION AND COOLING, COMBUSTION CHAMBER AND TURBOMACHINE WHILE ENHANCED|
    FR2929690B1|2008-04-03|2012-08-17|Snecma Propulsion Solide|COMBUSTION CHAMBER SECTORIZED IN CMC FOR GAS TURBINE|
    FR2930628B1|2008-04-24|2010-04-30|Snecma|ANNULAR COMBUSTION CHAMBER FOR TURBOMACHINE|US20140216044A1|2012-12-17|2014-08-07|United Technologoes Corporation|Gas turbine engine combustor heat shield with increased film cooling effectiveness|
    US10260748B2|2012-12-21|2019-04-16|United Technologies Corporation|Gas turbine engine combustor with tailored temperature profile|
    WO2014160299A1|2013-03-14|2014-10-02|United Technologies Corporation|Combustor panel with increased durability|
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    EP3071813B8|2013-11-21|2021-04-07|Raytheon Technologies Corporation|Axisymmetric offset of three-dimensional contoured endwalls|
    EP3084303B1|2013-12-19|2022-01-26|Raytheon Technologies Corporation|Thermal mechanical dimple array for a combustor wall assembly|
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    US10451280B2|2015-02-16|2019-10-22|United Technologies Corporation|Combustor panel having material transition region|
    FR3035707B1|2015-04-29|2019-11-01|Safran Aircraft Engines|COMBUSTION CHAMBER WITH TURBOMACHINE|
    US10222065B2|2016-02-25|2019-03-05|General Electric Company|Combustor assembly for a gas turbine engine|
    US10816202B2|2017-11-28|2020-10-27|General Electric Company|Combustor liner for a gas turbine engine and an associated method thereof|
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    法律状态:
    2019-01-08| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|
    2019-10-15| B06U| Preliminary requirement: requests with searches performed by other patent offices: procedure suspended [chapter 6.21 patent gazette]|
    2020-09-08| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|
    2020-12-08| B16A| Patent or certificate of addition of invention granted [chapter 16.1 patent gazette]|Free format text: PRAZO DE VALIDADE: 10 (DEZ) ANOS CONTADOS A PARTIR DE 08/12/2020, OBSERVADAS AS CONDICOES LEGAIS. |
    优先权:
    申请号 | 申请日 | 专利标题
    FR0906009A|FR2953907B1|2009-12-11|2009-12-11|COMBUSTION CHAMBER FOR TURBOMACHINE|
    FR0906009|2009-12-11|
    PCT/FR2010/052600|WO2011070273A1|2009-12-11|2010-12-02|Turbine engine combustion chamber|
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