• 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:
    Turbomachine combustion chamber annular wall, having between a cold side (16a, 18a) and a hot side (16b, 18b), a plurality of air intake ports (30) distributed in at least one circumferential row to allow air flowing from the cold side to enter the hot side, these air intake ports creating around them areas with a high thermal gradient, and a plurality of cooling orifices (32) to allow the air flowing on the cold side to enter the hot side to form a cooling air film along the annular wall, the annular wall further comprising, in areas of high thermal gradient, multi-hole holes; perforation (36) bent at an angle α greater than 90 °, this angle a being taken an input geometric axis Ae and a geometric output axis As of the multi-perforation hole, the geometric output axis of the hole of multi-perforation being inclined at an angle θ3 p relative to a normal N to the annular wall, in a so-called gyration direction at most perpendicular to the axial direction D of flow of the combustion gases.
    公开号:FR3037107A1
    申请号:FR1555050
    申请日:2015-06-03
    公开日:2016-12-09
    发明作者:Patrice Andre Commaret；Jacques Marcel Arthur Bunel；Romain Nicolas Lunel
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] BACKGROUND OF THE INVENTION The present invention relates to the general field of turbomachine combustion chambers. It is more particularly an annular wall for direct combustion chamber or reverse flow cooled by a so-called "multi-perforation" process. Typically, an annular turbomachine combustion chamber is formed of an inner annular wall (also called inner ferrule) and an outer annular wall (also called outer ferrule) which are connected upstream by a transverse wall forming the chamber bottom . The inner and outer shrouds are each provided with a plurality of holes and various air intake ports allowing air circulating around the combustion chamber to penetrate into the interior thereof. Thus, so-called "primary" and "dilution" holes are formed in these ferrules to convey air inside the combustion chamber. The air passing through the primary holes helps to create an air / fuel mixture that is burned in the chamber, while the air from the dilution holes is intended to promote the dilution of the same air / fuel mixture. The inner and outer shells are subjected to the high temperatures of the gases from the combustion of the air / fuel mixture. To ensure their cooling, additional holes called multi-perforation are also drilled through these ferrules over their entire surface. These multi-perforation orifices, generally inclined at 60 °, allow the air circulating outside the chamber to penetrate inside thereof by forming cooling air films along the shrouds. .
    [0002] However, in practice, it has been found that the zone of the inner and outer rings which is located around and in particular directly downstream of each of the primary or dilution holes, because of the laser drilling technology used, has an absence of orifices and thus benefits from a low level of cooling with the risk of formation and propagation of cracks that implies.
    [0003] In order to solve this problem, the Applicant has proposed in its application FR2982008 to carry out, just downstream of primary holes or dilution holes, additional cooling orifices arranged in a plane perpendicular to the direction of flow of the gases of FIG. combustion. However, these additional or so-called gyratory orifices (because at 90 °), if they make it possible to ensure effective cooling compared with the conventional multi-axial perforation where the air film is stopped by the presence of these holes, still present certain disadvantages in particular because they are powered by the single static pressure of the air 10 flowing around the combustion chamber. OBJECT AND SUMMARY OF THE INVENTION It is therefore an object of the present invention to overcome such drawbacks by proposing an annular wall of a combustion chamber which, by using total air pressure, provides a better cooling of the areas likely to see creep primers, especially those located directly downstream of the primary and dilution holes, appear. For this purpose, there is provided an annular wall of a turbomachine combustion chamber, comprising a cold side and a hot side, said annular wall comprising: a plurality of air intake ports distributed in at least one circumferential row to allow air flowing from said cold side to enter said hot side, said air intake ports creating around them areas with a high thermal gradient, and. a plurality of cooling ports for allowing air flowing from said cold side to enter said hot side to form a cooling air film along said annular wall, said cooling ports being distributed in a plurality of rows Circumferentially spaced axially from each other and the geometric axes of each of said cooling orifices being inclined, in an axial direction D of flue gas flow, by an angle of inclination 01 with respect to a normal N to said annular wall, said annular wall being characterized in that it further comprises, in said zones with a high thermal gradient, perforation holes bent at an angle α greater than 90 °; , said angle a being taken between an input geometric axis Ae and an output geometric axis As of said multi-perforation hole, said output geometric axis of said tr or multi-perforation being inclined at an angle 83 relative to said normal N to said annular wall, in a so-called gyration direction at most perpendicular to said axial direction D of flow of the combustion gases. The presence of these cooling orifices bent in the wall of the combustion chamber so that they are fed as multi-perforation holes on the cold side while opening on the hot side as gyratory multi-perforation holes, allows for optimal cooling with the best of classic axial mufti-perforation as well as multi-perforation gyratory. Advantageously, said angle α is between 90 ° and 170 ° and said direction of gyration is inclined with respect to said axial direction D of flue gas flow at an angle of between 50 ° and 90 °. Preferably, said multi-perforation holes have a diameter d3, preferably identical to said diameter d1 of said cooling orifices and said outlet inclination 83 is preferably identical to said inclination 81 of said cooling orifices, each of said multi-perforation holes. perforation may have a variable profile, so as to optimize the cooling locally. Advantageously, said input geometrical axis of said multi-perforation hole is inclined at an angle 84 with respect to said normal N to said annular wall, in an axial direction D of flow of the combustion gases, said input inclination 84 being preferably identical to said inclination 81 of said cooling orifices. Preferably, said multi-perforation holes divide in two from a bent portion to form two air outlets opening on the hot side and the or both air outlets opening on the hot side of said multi-perforation holes can or may have a variable diameter d3 cone-shaped. According to the contemplated embodiment, said air intake orifices are primary holes allowing air flowing from said cold side to enter said hot side to create an air / fuel mixture or said air ports. air intake are dilution holes allowing air flowing from said cold side to enter said hot side to ensure dilution of the air / fuel mixture. The present invention also relates to a combustion chamber and a turbomachine (having a combustion chamber) comprising an annular wall as defined above. BRIEF DESCRIPTION OF THE DRAWINGS Other features and advantages of the present invention will be apparent from the description given below, with reference to the accompanying drawings which illustrate an embodiment thereof devoid of any limiting character. In the figures: - Figure 1 is a longitudinal sectional view of a turbomachine combustion chamber in its environment; FIG. 2 is a partial view in developed of one of the annular walls of the combustion chamber of FIG. 1 according to one embodiment of the invention; - Figures 3 and 4 are partial views respectively from above and in perspective of a portion of the annular wall of Figure 2; and FIGS. 5 and 6 show two alternative embodiments of the multi-perforation holes bent in one of the annular walls of the combustion chamber of FIG. 1. DETAILED DESCRIPTION OF THE INVENTION FIG. environment a combustion chamber 10 for turbomachine. Such a turbomachine comprises in particular a compression section (not shown) in which air is compressed before being injected into a chamber housing 12, then into the combustion chamber 10 mounted inside thereof.
    [0004] The compressed air is introduced via air intake ports into the combustion chamber and mixed with fuel prior to being burned. The gases resulting from this combustion are then directed to a high-pressure turbine 14 disposed at the outlet of the combustion chamber. The combustion chamber is of the annular type. It is formed of an inner annular wall 16 and an outer annular wall 18 which are joined upstream by a transverse wall 20 forming the chamber bottom. It can be direct as illustrated or reverse flow. In this case, a return elbow is placed between the combustion chamber and the turbine distributor. The inner annular walls 16 and outer 18 extend along a longitudinal axis slightly inclined relative to the longitudinal axis 22 of the turbomachine. The chamber bottom 20 is provided with a plurality of openings 20A in which are mounted a same plurality of fuel injectors 24. The chamber housing 12, which is formed of an inner envelope 12a and an envelope external 12b, household with the combustion chamber 10 of the annular spaces 26 in which is admitted compressed air for combustion, dilution and cooling of the chamber. The inner annular walls 16 and outer 18 each have a cold side 16a, 18a disposed on the side of the annular space 26 in which the compressed air flows and a hot side 16b, 18b facing the inside of the combustion chamber. (see Figure 4). The combustion chamber 10 is divided into a so-called "primary" zone (or combustion zone) and a so-called "secondary" zone (or dilution zone) situated downstream from the previous one (the downstream means with respect to an axial general direction of flow of the gases resulting from the combustion of the air / fuel mixture inside the combustion chamber and represented by the arrow D). The air that feeds the primary zone of the combustion chamber 25 is introduced by a circumferential row of primary holes 28 formed in the inner annular walls 16 and outer 18 of the chamber over the entire circumference of these annular walls. These primary holes have a downstream edge aligned on the same line 28A. As for the air supplying the secondary zone of the chamber, it borrows a plurality of dilution holes 30 also formed in the inner annular walls 16 and outer 18 all around the circumference of these annular walls. These dilution holes 30 are aligned in a circumferential row which is axially offset downstream from the rows of the primary holes 28 and may have different diameters with in particular alternating large and small holes. In the configuration illustrated in FIG. 2, however, these dilution holes of different diameters have a downstream edge aligned on the same line 30A. In order to cool the inner and outer annular walls 16 of the combustion chamber which are subjected to the high temperatures of the combustion gases, a plurality of cooling orifices 32 (illustrated in FIGS. 2 to 4) are provided. In the case of an inverted flow chamber, the return elbow is also provided with such orifices. These orifices 32, which provide cooling of the walls 16, 10 18 by multiple holes also called "multi-perforation", are distributed in a plurality of circumferential rows spaced axially from each other. These rows of multi-perforation orifices generally cover the entire surface of the annular walls of the combustion chamber. The number and diameter of the cooling orifices 32 are identical in each of the rows. The pitch pl between two orifices of the same row is constant and may be identical or not for all the rows. Furthermore, the adjacent rows of cooling orifices are arranged so that the orifices 32 are staggered as shown in FIG. 2.
    [0005] As illustrated in FIG. 4, the cooling orifices 32 generally have an angle of inclination 81 with respect to a normal N to the annular wall 16, 18 through which they are pierced. This inclination 81 allows the air passing through these orifices to form a film of air along the hot side 16b, 18b of the annular wall.
    [0006] With respect to non-inclined orifices, it makes it possible to increase the surface of the annular wall which is cooled. In addition, the inclination 01 of the cooling orifices 32 is directed so that the air film thus formed flows in the direction of flow of the combustion gases inside the chamber (shown diagrammatically by the arrow D ).
    [0007] By way of example, for an annular wall 16, 18 made of metal or ceramic material and having a thickness of between 0.6 and 3.5 mm, the diameter di of the cooling orifices 32 may be between 0.3 and 1 mm (preferentially between 0.4 and 0.6 mm), the pi pitch between 1 and 10 mm and their inclination 01 between 35 + 30 ° and + 70 °, typically + 60 °. By way of comparison, for an annular wall having the same characteristics, the primary holes 28 and the dilution holes 30 have a diameter of the order of 4 to 20 mm. In addition, to provide efficient cooling downstream of the dilution holes, each annular wall 16, 18 of the combustion chamber may further comprise disposed directly downstream of these dilution holes 30 (but a similar configuration downstream of the primary holes 28 is also possible to limit the rise of the thermal gradient at these holes and thus avoid the formation of cracks) and distributed along several circumferential rows from the upstream transition axis 30A, a plurality of additional holes of However, unlike the preceding cooling orifices which deliver a film of air flowing in the axial direction D, the air film delivered by these additional orifices flows in a perpendicular direction because of their arrangement. at 90 ° in a plane perpendicular to this axial direction D of flue gas flow. This multi-perforation made perpendicularly to the axis of the turbomachine (in the following description, it will speak of multiperforation gyratory as opposed to the axial multi-perforation of the cooling orifices) allows to bring these additional holes 20 dilution holes (or primary holes if necessary) and therefore limit the rise of the thermal gradient at these holes. The additional orifices 34 of the same row have the same diameter d2, preferably identical to the diameter di of the cooling orifices 32, are spaced by a constant pitch p2 which may or may not be identical to the pitch p1 between the cooling orifices. 32 and have an inclination 02, preferably identical to the inclination 01 of the cooling orifices 32 but arranged in a perpendicular plane. However, these characteristics of the additional orifices 34 may, while remaining within the previously defined ranges of values, be substantially different from those of the cooling orifices 32, that is to say that the inclination of the additional orifices the same row with respect to a normal N to the annular wall 16, 18 may be different from that 01 of the cooling orifices, and the diameter d2 of the additional orifices of the same row may be different from that of the cooling orifices 32.
    [0008] According to the invention, optimum cooling around the dilution holes as primary holes is obtained by using the dynamic pressure on the cold side (between the casing and the wall of the chamber) while retaining the effect of the hot side gyration. To do this, provision is made in the zones with a high thermal gradient such as the primary or dilution holes, multi-perforation holes 36 bent at an angle α greater than 90 ° in the wall of the chamber, so that they are fed by the total pressure of the air cold side, as for an axial multi-perforation, while opening on the hot side as 10 in multi-perforation gyratory. Indeed, as is known, in the case of the 90 ° multi-perforation, the holes are powered by the static pressure of the air between the housing 12 and the chamber wall while in the case of the multi -Perforation axial holes are powered by the total air pressure. Being fed with static pressure means that the dynamic air pressure (total P - static P) is not used between the housing and the chamber wall. However, this dynamic pressure is even greater at the primary holes and dilution holes. The angle typically between 90 ° and 170 ° is taken between an input geometric axis Ae and an output geometric axis As of the multi-perforation hole, the output geometric axis (axis of the hot-wall side hole). ) being inclined at an angle 03 with respect to the normal N to the annular wall but in a plane itself inclined by an angle [3 in a so-called gyration direction at most perpendicular to the axial direction D 25 of flow combustion gases. Typically this angle of turn 13 is between 50 ° and 90 °. These multi-perforation holes 36 have a diameter d3, preferably identical to the diameter d1 of the cooling orifices 32 and the inclination at the outlet 83 is preferably identical to the inclination 81 of the cooling orifices 32. Likewise, input geometric axis (cold side hole axis) may be straight (parallel to the normal N) or preferably have an inclination 84 (preferably identical to the inclination 81 of the cooling orifices 32) with respect to this normal in the axial direction D of flue gas flow.
    [0009] However, these various characteristics may, while remaining within the previously defined ranges of values for the cooling orifices 32, be substantially different. It may be noted that in these zones with a high thermal gradient, the profile of each multi-perforation hole 36 can advantageously be variable, that is to say with a diameter d 3, inclinations 83 and 84 (therefore an angle a) and a gyration angle f3 different from one hole to another, so as to optimize the cooling locally. It should also be noted that for a 0.4 mm diameter hole, the lateral exchange surface (air-wetted surface passing through the hole) is identical for the 3 types of multi-perforation holes. Thus, this lateral exchange surface S can be determined by the following formula: S = n * d3 * L (with L the length of the hole) Either for a straight axial hole 32 equal to 3 mm 2, or for a hole 34 giratory right at 90 ° S equal to 3.04 mm2, and 15 or for a hole 36 bent at 90 ° S equal to 3.06 mm2. The bent hole does not degrade the cooling of the wall by the forced convection of the air flowing through it. With the multi-perforation composed of bent holes, the use of the dynamic pressure of the air flowing between the casing and the wall of the combustion chamber which was not used with the multi-perforation, allows to cool simply and much more efficiently around the primary holes and dilution by judicious placement of these bent holes. According to an alternative embodiment illustrated in FIG. 5, the air outlet 25 of the multi-perforation hole 36 opening on the hot side can have a variable diameter d3 in the shape of a cone and therefore no longer be constant as before but with the diameter d3. increasing as it gets closer to this air outlet. Likewise, as shown in FIG. 6, the air outlet opening on the hot side can, after the bent portion, be divided into two parts so as to increase the exchange surface between the cooling air and the wall. The diameter d3 will preferably be identical throughout the multi-perforation hole 36, but the angles α1 and α2 between the input geometrical axis and the respective output geometrical axes of the two parts after the bend may be identical as illustrated. or even different. However, as previously, the air outlets forming the end portion of the hole may also be of variable diameter.
    权利要求:
    Claims (12)
    [0001]
    REVENDICATIONS1. Turbomachine combustion chamber annular wall (16, 18) having between a cold side (16a, 18a) and a hot side (16b, 18b):. a plurality of air intake ports (28, 30) distributed in at least one circumferential row to allow air flowing from said cold side to enter said hot side, said air intake ports creating around from them zones with a high thermal gradient, and. a plurality of cooling orifices (32) for allowing air flowing from said cold side to enter said hot side to form a cooling air film along said annular wall, said cooling orifices being distributed according to a a plurality of circumferentially spaced rows spaced axially from each other and the geometric axes of each of said cooling orifices being inclined, in an axial direction D of flow of the combustion gases, with an inclination angle θ1 with respect to a normal N. to said annular wall, said annular wall being characterized in that it further comprises, in said zones with a high thermal gradient, multi-perforation holes (36) bent at an angle α greater than 90 °, said angle a being taken between an input geometric axis Ae and an output geometric axis As of said multi-perforation hole, said output geometric axis of said multi-hole ti-perforation being inclined at an angle 03 with respect to said normal N to said annular wall, in a so-called gyration direction at most perpendicular to said axial direction D flue gas flow.
    [0002]
    2. Wall according to claim 1, characterized in that said angle a is between 90 ° and 170 °.
    [0003]
    3. Wall according to claim 1, characterized in that said direction of gyration is inclined with respect to said axial direction D of flue gas flow at an angle of between 50 ° and 90 °. 35 3037107 12
    [0004]
    4. Wall according to any one of claims 1 to 3, characterized in that said multi-perforation holes have a diameter d3, preferably identical to said diameter dl of said cooling orifices and said outlet inclination 03 is preferably identical to said inclination 01 of said cooling orifices.
    [0005]
    5. Wall according to claim 4, characterized in that said input geometrical axis of said multi-perforation hole is inclined at an angle 84 with respect to said normal N to said annular wall, in an axial direction D of flow of the combustion gases, said inlet inclination 04 being preferably identical to said inclination 81 of said cooling orifices.
    [0006]
    6. Wall according to claim 4 or claim 5, characterized in that each of said multi-perforation holes has a variable profile, so as to optimize the cooling locally.
    [0007]
    7. Wall according to any one of claims 1 to 6, characterized in that said multi-perforation holes divide in two from a bent portion to form two air outlets opening on the hot side.
    [0008]
    8. Wall according to any one of claims 1 to 7, characterized in that the or two air outlets opening on the hot side of said multi-perforation holes has (s) a diameter d3 variable cone-shaped.
    [0009]
    9. Wall according to claim 1, characterized in that said air intake orifices are primary holes (28) allowing air flowing from said cold side to penetrate said hot side to create an air / fuel mixture .
    [0010]
    10. Wall according to claim 9, characterized in that said air intake orifices are dilution holes (30) allowing air flowing from said cold side to penetrate said hot side to ensure dilution of the air / fuel mixture. 25 30 35 3037107 13
    [0011]
    11. A turbomachine combustion chamber (10) comprising at least one annular wall (16, 18) according to any one of claims 1 to 10. 5
    [0012]
    12. A turbomachine comprising a combustion chamber (10) having at least one annular wall (16, 18) according to any one of claims 1 to 10.
    类似技术:
    公开号 | 公开日 | 专利标题
    EP3267111B1|2022-02-16|Annular wall of a combustion chamber with improved cooling at the primary and/or dilution holes
    EP3303774B1|2019-05-01|Annular wall of a combustion chamber with optimised cooling
    EP1777458B1|2015-08-12|Improvement of the performances of a combustion chamber by multiperforating the walls
    CA2639980C|2016-04-12|Turbomachine combustion chamber
    WO2016146932A1|2016-09-22|Turbine ring assembly made from ceramic matrix composite material
    EP2510284B1|2018-08-29|Turbine engine combustion chamber
    EP1593814A1|2005-11-09|Feather seal configuration for a gas turbine shroud
    EP1489359A1|2004-12-22|Annular combustion chamber for turbomachine
    WO2015079166A1|2015-06-04|Combustion assembly having facilitated access to the prevaporization tubes
    FR2975465A1|2012-11-23|WALL FOR TURBOMACHINE COMBUSTION CHAMBER COMPRISING AN OPTIMIZED AIR INLET ORIFICE ARRANGEMENT
    EP3318725A1|2018-05-09|Connection assembly for cooling a turbine of a turbine engine
    FR2982009A1|2013-05-03|Annular wall for combustion chamber of turboshaft engine, has dilution holes whose rectangular sections exhibit equal axial length and width in direction transverse to axial direction of flow of combustion gases
    FR2979416A1|2013-03-01|WALL OF COMBUSTION CHAMBER
    FR3072448B1|2019-10-18|TURBOMACHINE COMBUSTION CHAMBER
    FR3017928B1|2019-07-26|TURBOMACHINE WITH EXTERNAL FLANGE OF "SANDWICH" COMBUSTION CHAMBER
    FR3021351B1|2019-07-12|TURBOMACHINE WALL HAVING AT LEAST ONE PORTION OF COOLING ORIFICES OBSTRUCTIONS
    FR3019270A1|2015-10-02|ANNULAR ROOM OF COMBUSTION CHAMBER HAVING IMPROVED COOLING BODIES AT FLANGE JOINT LEVELS
    FR2999277A1|2014-06-13|Annular internal or external wall for e.g. direct flow combustion chamber, of turboshaft engine, has cooling holes whose drilling axes are directed according to direction of air flow so as to maintain supply of air axially across holes
    WO2013045802A1|2013-04-04|Turbine engine combustion chamber
    EP3569929A1|2019-11-20|Assembly for a turbine engine combustion chamber
    FR3098569A1|2021-01-15|ANNULAR WALL FOR TURBOMACHINE COMBUSTION CHAMBER INCLUDING PRIMARY HOLES, DILUTION HOLES AND INCLINED COOLING PORTS
    FR3085743A1|2020-03-13|ANNULAR COMBUSTION CHAMBER FOR A TURBOMACHINE
    FR3033028A1|2016-08-26|COMBUSTION CHAMBER OF TURBOMACHINE HAVING A PENETRANT PART WITH OPENING
    FR2920221A1|2009-02-27|Combustion chamber for turbine engine, has deflector extending annular wall towards interior of chamber and downstream, and diverting air from exterior towards interior and downstream of chamber
    FR3101915A1|2021-04-16|Turbomachine turbine ring comprising internal cooling pipes
    同族专利:
    公开号 | 公开日
    CA2987526A1|2016-12-08|
    CN107683391B|2020-02-18|
    EP3303774B1|2019-05-01|
    US10760436B2|2020-09-01|
    WO2016193589A1|2016-12-08|
    EP3303774A1|2018-04-11|
    FR3037107B1|2019-11-15|
    RU2017145251A|2019-07-12|
    RU2017145251A3|2019-10-01|
    ES2729098T3|2019-10-30|
    CN107683391A|2018-02-09|
    US20180142563A1|2018-05-24|
    RU2718371C2|2020-04-02|
    BR112017025792A2|2018-08-07|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    WO2013060987A2|2011-10-26|2013-05-02|Snecma|Annular wall of a combustion chamber with improved cooling at the primary and/or dilution holes|
    EP2759772A1|2013-01-23|2014-07-30|Honeywell International Inc.|Combustors with complex shaped effusion holes|
    US20150107798A1|2013-10-18|2015-04-23|Rolls-Royce Deutschland Ltd & Co Kg|Unknown|
    SU1176678A1|1984-03-01|1995-08-27|О.Г. Жирицкий|Fire tube of internal combustion gas-turbine engine|
    CN85107191A|1984-10-04|1986-09-24|西屋电气公司|Impact type cooling gas turbine firing chamber with interior air film cooling|
    SU1454021A1|1987-03-19|1995-07-20|А.С. Косой|Flame tube for combustion chamber of gas-turbine engine|
    US4923371A|1988-04-01|1990-05-08|General Electric Company|Wall having cooling passage|
    US5181379A|1990-11-15|1993-01-26|General Electric Company|Gas turbine engine multi-hole film cooled combustor liner and method of manufacture|
    US5241827A|1991-05-03|1993-09-07|General Electric Company|Multi-hole film cooled combuster linear with differential cooling|
    US5307637A|1992-07-09|1994-05-03|General Electric Company|Angled multi-hole film cooled single wall combustor dome plate|
    US5419681A|1993-01-25|1995-05-30|General Electric Company|Film cooled wall|
    RU2066424C1|1994-04-28|1996-09-10|Акционерное общество "Авиадвигатель"|Combustion chamber of gas-turbine engine|
    US6383602B1|1996-12-23|2002-05-07|General Electric Company|Method for improving the cooling effectiveness of a gaseous coolant stream which flows through a substrate, and related articles of manufacture|
    GB9803291D0|1998-02-18|1998-04-08|Chapman H C|Combustion apparatus|
    US6627019B2|2000-12-18|2003-09-30|David C. Jarmon|Process for making ceramic matrix composite parts with cooling channels|
    RU2260748C2|2003-12-02|2005-09-20|Открытое акционерное общество "Авиадвигатель"|Combustion chamber for gas-turbine engine|
    US7131818B2|2004-11-02|2006-11-07|United Technologies Corporation|Airfoil with three-pass serpentine cooling channel and microcircuit|
    CA2627958C|2005-11-01|2011-03-22|Ihi Corporation|Turbine component|
    US7856830B2|2006-05-26|2010-12-28|Pratt & Whitney Canada Corp.|Noise reducing combustor|
    US7942006B2|2007-03-26|2011-05-17|Honeywell International Inc.|Combustors and combustion systems for gas turbine engines|
    US8146364B2|2007-09-14|2012-04-03|Siemens Energy, Inc.|Non-rectangular resonator devices providing enhanced liner cooling for combustion chamber|
    US8092176B2|2008-09-16|2012-01-10|Siemens Energy, Inc.|Turbine airfoil cooling system with curved diffusion film cooling hole|
    FR2946413B1|2009-06-04|2011-07-15|Snecma|GAS TURBINE ENGINE COMBUSTION CHAMBER WITH MULTI-PERFORATED WALL ELEMENT|
    JP5653705B2|2010-09-30|2015-01-14|三菱重工業株式会社|Recovery air cooling gas turbine combustor cooling structure|
    DE102010051638A1|2010-11-17|2012-05-24|Rolls-Royce Deutschland Ltd & Co Kg|Gas turbine combustor with a cooling air supply device|
    FR2982009B1|2011-10-26|2013-12-13|Snecma|ANNULAR ROOM OF COMBUSTION CHAMBER WITH IMPROVED COOLING AT PRIMARY HOLES AND / OR DILUTION HOLES|
    US20160245094A1|2015-02-24|2016-08-25|General Electric Company|Engine component|US11015529B2|2016-12-23|2021-05-25|General Electric Company|Feature based cooling using in wall contoured cooling passage|
    CN108870445A|2018-04-11|2018-11-23|西北工业大学|A kind of flame combustion chamber tube wall surface using the Y shape effusion wall type of cooling|
    US10801727B2|2018-07-06|2020-10-13|Rolls-Royce North American Technologies Inc.|System for combustor cooling and trim air profile control|
    US11029027B2|2018-10-03|2021-06-08|Raytheon Technologies Corporation|Dilution/effusion hole pattern for thick combustor panels|
    DE102019105442A1|2019-03-04|2020-09-10|Rolls-Royce Deutschland Ltd & Co Kg|Method for producing an engine component with a cooling duct arrangement and engine component|
    US20210018173A1|2019-07-16|2021-01-21|United Technologies Corporation|Effusion cooling for dilution/quench hole edges in combustor liner panels|
    法律状态:
    2016-06-08| PLFP| Fee payment|Year of fee payment: 2 |
    2016-12-09| PLSC| Search report ready|Effective date: 20161209 |
    2017-04-26| PLFP| Fee payment|Year of fee payment: 3 |
    2018-06-05| PLFP| Fee payment|Year of fee payment: 4 |
    2018-09-14| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20180809 |
    2019-05-22| PLFP| Fee payment|Year of fee payment: 5 |
    2020-05-20| PLFP| Fee payment|Year of fee payment: 6 |
    2021-05-19| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1555050A|FR3037107B1|2015-06-03|2015-06-03|ANNULAR ROOM OF COMBUSTION CHAMBER WITH OPTIMIZED COOLING|
    FR1555050|2015-06-03|FR1555050A| FR3037107B1|2015-06-03|2015-06-03|ANNULAR ROOM OF COMBUSTION CHAMBER WITH OPTIMIZED COOLING|
    BR112017025792-0A| BR112017025792A2|2015-06-03|2016-05-27|annular wall of turbomachine combustion chamber, turbomachine combustion chamber, and turbomachine.|
    US15/579,006| US10760436B2|2015-06-03|2016-05-27|Annular wall of a combustion chamber with optimised cooling|
    ES16733638T| ES2729098T3|2015-06-03|2016-05-27|Annular wall of a combustion chamber with optimized cooling|
    PCT/FR2016/051263| WO2016193589A1|2015-06-03|2016-05-27|Annular wall of a combustion chamber with optimised cooling|
    CN201680032599.4A| CN107683391B|2015-06-03|2016-05-27|Annular wall of a combustion chamber with optimized cooling|
    RU2017145251A| RU2718371C2|2015-06-03|2016-05-27|Annular wall of combustion chamber with optimized cooling|
    EP16733638.7A| EP3303774B1|2015-06-03|2016-05-27|Annular wall of a combustion chamber with optimised cooling|
    CA2987526A| CA2987526A1|2015-06-03|2016-05-27|Annular wall of a combustion chamber with optimised cooling|
    [返回顶部]