• 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:
    An aircraft turbine engine stator (40) having an annular array of stationary blades (44) and an annular arm array (48), characterized in that the trailing edges (42) of the stationary vanes are located substantially in a first transverse plane (P1) located downstream of a second transverse plane (P2) passing substantially through the leading edges (46) of the arms.
    公开号:FR3027053A1
    申请号:FR1459721
    申请日:2014-10-10
    公开日:2016-04-15
    发明作者:Simon Pierre Claude Charbonnier；Matthieu Yoann Perrier
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention relates to an aircraft turbomachine stator, and in particular to a stator comprising at least one annular row of stationary vanes and an annular row of arms. STATE OF THE ART In general, an aircraft turbine engine comprises, upstream and downstream, in the direction of flow of the gases, a fan, at least one compressor, an annular combustion chamber, at least one turbine, and a combustion gas ejection nozzle. In the case of a turbofan engine, the flow of air through the fan splits into a primary flow that feeds the engine and a secondary flow that flows around the engine.
    [0002] In a conventional manner, the motor comprises at least one stator and at least one rotor. In the case of a multi-body turbojet, the engine may for example comprise two rotors or bodies, one low pressure and the other high pressure. The turbojet can thus comprise a low pressure body having a first shaft connecting a low pressure compressor to a low pressure turbine, and a high pressure body having a second shaft connecting a high pressure compressor to a high pressure turbine. The air entering the engine is successively compressed in the low pressure compressor and the high pressure compressor before being mixed with fuel which is burned in the combustion chamber. The combustion gases then relax in the high pressure turbine and the low pressure turbine to drive the low pressure shaft in rotation which in turn drives the fan shaft. The stator of the turbomachine comprises annular structural housings 30, that is to say casings sufficiently rigid to transmit forces. A turbomachine includes in particular an inlet casing 3027053 2 extends downstream of the blower and an intermediate casing which extends between the low and high pressure compressors. Each of these stator housings generally comprises generally structural arms which extend substantially radially between two annular walls, respectively internal and external, these arms being tubular for the passage of servitudes from the inside of the inner wall up to the outside of the outer wall. In order to reduce the consumption of the engines, some turbomachines include gearboxes. The stators of these turbomachines generally comprise an annular row of vanes upstream of the structural arms of the casing and an annular row of vanes with variable spacing downstream of the structural arms. The intermediate casing of this type of turbomachine, which is located downstream of the low pressure compressor, thus comprises an annular row of structural arms also called primary arms interposed between an annular row of blades (called blades IGV, acronym for English Inlet Guide Vane) and an annular row of variable pitch vanes (called VSV blades, stands for Variable Stator Vane). The inlet casing, which is located upstream of the low pressure compressor, also comprises an annular row of structural arms interposed between an annular row of stationary blades and an annular row of variable pitch vanes. This type of stator with three successive rows of blades and arms (two rows of blades and one row of arms) is very penalizing because it has a negative impact on the mass and the pressure losses due to an axial space requirement. important. The present invention provides a simple, effective and economical solution to this problem. SUMMARY OF THE INVENTION The invention proposes an aircraft turbomachine stator comprising an annular row of stationary blades and an annular row of arms, for example structural, said stator having an axis of revolution, characterized in that the trailing edges of the vanes are located substantially in a first plane transverse to said axis and located downstream of a second plane transverse to said axis and passing substantially through the leading edges of the arms.
    [0003] The invention makes it possible to reduce the axial space requirement of a stator comprising an annular row of stationary blades and an annular row of arms, by at least partially axially interleaving these rows. Thus, in contrast to the prior art in which the vanes were upstream of the arms, the vanes are here positioned at least partly between the arms. The axial size of the stator can thus be significantly reduced compared with those of the prior art, which has a positive impact on the mass and the pressure drops at the stator crossing. Although the rows of stationary blades and arms are arranged in a particular way according to the invention, the parameters and the aerodynamic properties of the profiles of the blades and the arms are advantageously preserved. The arms thus preferably retain their thicknesses, their shapes and their secondary functions of passage of easements. The blades can retain their function as a rectifier of the airflow 20 (exiting for example from the fan of the turbomachine). This makes it possible to keep the triangles of the speeds upstream and downstream of the stator. The leading edges of the blades can be located in a third plane transverse to said axis and located upstream of the second transverse plane.
    [0004] This improves the aerodynamic performance of the stator in the different phases of flight. Indeed, the fact of shifting the leading edges of the blades upstream relative to those of the arms makes it possible to promote the passage of the flow in the blades. This furthermore makes it possible to soften the deflection of the outlet flow of the blower by avoiding the direct deflection by the arms. Thus, there is less pressure drop and the efficiency of the turbomachine is improved.
    [0005] In the present application, transverse plane means a plane substantially perpendicular to the longitudinal axis or revolution of the stator, which is generally the longitudinal axis of the turbomachine. Advantageously, the median longitudinal planes of the arms are inclined relative to the longitudinal axis of the stator. The median longitudinal planes of at least some of the arms and at least some of the fixed blades are preferably substantially parallel or slightly inclined. This facilitates the nesting of the row of blades in the row of arms. The inclination or wedging of the arms makes it possible to further reduce the axial size of the stator without decreasing the actual rope of the arms, and thus without modifying their thickness (it is thus possible to preserve the thickness / rope ratio as well as the number and distribution of the arm compared to the prior art). The stator may further comprise an annular array of variable pitch vanes located directly downstream of the annular row of arms. As in the aforementioned case, the stator according to the invention thus comprises two rows of blades and a row of arms. The annular row of variable pitch vanes may comprise first variable pitch vanes which are located substantially in the extension of the arms, and second variable pitch vanes which are located between the first variable pitch vanes. Advantageously, the first variable-pitch blades have aerodynamic profiles different from those of the second variable-pitch blades. The fact of inserting the vanes between the arms has an impact on the aerodynamic performance of the stator because the velocity triangle is not constant azimuthally downstream of the arms. To overcome this drawback, the variable-pitch vanes located downstream of the arms have profiles that are not all identical. Two different profiles are used depending on whether the blade is directly downstream of an arm or downstream of a fixed blade.
    [0006] The leading edges of the first blades with variable pitch are preferably located as close as possible to the trailing edges of the arms so as to further reduce the axial stator space requirement, but also so that these first blades act as "shutters". To ensure the operability of the low and high pressure compressors under all flight conditions. The stator may comprise two annular walls respectively internal and external, between which the rows of blades and arms extend.
    [0007] Each variable-pitch blade may comprise at its outer radial end a cylindrical pivot which is mounted in a chimney of the outer wall. The outer wall may comprise, upstream of the arms, an annular row of through-slots of discharge air passage and means for shutting, preferably adjustable, of these slots. The slots and the closure means form a discharge valve also called VBV valve, an acronym for Variable Bleed Valve. As a variant, the outer wall comprises, between arms, at least one through opening of the discharge air passage and at least one shutter door, preferably adjustable, of this opening. In a particular embodiment of the invention, an estimate of the mass reduction provided by the invention would be 5 to 10% of the mass of the module per casing of this type. This is mainly due to the reduction in axial length (and therefore the reduction of the internal and external walls of housings in the primary and secondary streams), related to the integration of the vanes between the arms, the reduction of the games axial (especially between the rows of arms and blades with variable setting), and the setting of the arms (which can be between 20 and 30 ° - which allows to gain about 10% of rope - against 0 ° in the prior art).
    [0008] 3027053 6 An additional, non-quantifiable gain is the reduction in mass of the pipes and harnesses that run around the fan casing, by reducing the distances traveled. The present invention also relates to an aircraft turbine engine 5, such as a turbojet engine or a turboprop engine, characterized in that it comprises at least a stator as described above. In the case where the turbomachine comprises a blower and low and high pressure compressors, the stator can be mounted between the blower and the low pressure compressor and / or between the low and high pressure compressors. Alternatively, when the turboprop comprises for example two contrarotating external propellers located in the vicinity of the downstream end of the turboprop engine with reference to the flow of gases in its engine, the stator can be mounted at the inlet of the low pressure compressor.
    [0009] DESCRIPTION OF THE FIGURES The invention will be better understood and other details, characteristics and advantages of the invention will become apparent on reading the following description given by way of nonlimiting example and with reference to the appended drawings in which: FIG. 1 is a partial schematic half-view in axial section of an aircraft turbomachine according to the prior art, seen from the side; FIG. 2 is a very diagrammatic view of part of the turbomachine of FIG. 3 is a partial diagrammatic view in axial section of an aircraft turbomachine according to the invention, seen from the side - FIG. 4 is a very schematic view of a part of the turbomachine of FIG. 3, top view, - Figure 5 is a view corresponding to Figure 3 and showing an embodiment of an inlet housing according to the invention, and 3027053 7 - Figures 6 and 7 are views corresponding to Figure 3 and represented design variants of an intermediate casing according to the invention. DETAILED DESCRIPTION Reference is first made to FIG. 1 which represents a turbomachine 10 according to the prior art, this turbomachine 10 being here a turbofan engine. The invention is described with respect to this example but it is clear that the invention can be applied to other turbomachine architectures.
    [0010] The turbomachine 10 comprises, from upstream to downstream, in the direction of flow of the gases, a fan 12 which generates a flow which splits into two coaxial flows, the primary flow supplying the engine which comprises a low pressure compressor 14, a high pressure compressor (not shown), a combustion chamber (not shown), high and low pressure turbines (not shown) and a combustion gas ejection nozzle (not shown). These motor modules (blower, compressors, combustion chamber, turbines) are surrounded by annular structural stator housings. The turbomachine 10 thus comprises a plurality of successive annular casings 20 among which an inlet casing 18 upstream of the low pressure compressor 14, and an intermediate casing 20 between the low and high pressure compressors. The inlet housing 18 includes an annular row of structural arms 24 (or primary arms) which is interposed between an annular row of stationary vanes 22 (or IGV vanes) and an annular array of variable pitch vanes 26 (or VSV blades). As is also visible in FIG. 2, the row of stationary vanes 22 is located between the fan 12 and the row of arms 24, and the row of variable valve vanes 26 is located between the row of arms 24 and a row of vanes 24. mobile rotor wheel 28 of the low pressure compressor 14.
    [0011] In the same way, the intermediate casing 20 comprises an annular row of structural arms (or primary arms) which is interposed between an annular row of stationary vanes (or vanes IGV) and an annular row of variable-pitch vanes ( or VSV blades). The stationary blade row is located between a rotor wheel of the low pressure compressor and the row of arms, and the row of variable pitch vanes is located between the row of arms and a rotor wheel of the high pressure compressor. The reference 34 in FIG. 2 represents the velocity triangles of the air flow entering the fan 12, and the references 36, 38 represent the velocity triangles of the primary air flow upstream of the stationary vanes 22 and downstream of the Variable valve blades 26. As explained above, each of the inlet casing 18 and intermediate 20 forms, axially with the rows of blades 22, 26, a bulky stator.
    [0012] The invention overcomes this problem by the axial interlocking of the row of vanes in the stator arm row. Figures 3 and 4 show an embodiment of a stator or inlet housing according to the invention, this embodiment can naturally be applied to an intermediate casing.
    [0013] P1 denotes a transverse plane (perpendicular to the longitudinal axis A or revolution of the stator 40 and the turbomachine) substantially passing through the trailing edges 42 of the vanes 44, P2 a transverse plane passing substantially through the edges of Attack 46 of the arms 48, P3 a transverse plane passing substantially through the leading edges 50 of the vanes 44, P4 a transverse plane passing substantially through the trailing edges 52 of the arms 48, and P5 a transverse plane passing substantially through the leading edges 54 of variable pitch vanes 56, 58. In the example shown, P1 is downstream of P2 which is downstream of P3. Moreover, P5 is downstream from P4 and at a short axial distance from it.
    [0014] The variable-pitch vanes 56, 58 are therefore located directly downstream of the arms 48. In addition, H1 denotes a median longitudinal plane for each arm 48 and H2 a median longitudinal plane for each fixed blade 44.
    [0015] The planes H1 and H2 are here inclined with respect to the axis A and are substantially parallel to each other. As opposed to the prior art, the arms 48 therefore do not have an axial orientation but are on the contrary "wedged" (at an angle of 20-30 ° for example) around a substantially radial axis relative to the axis A.
    [0016] Each arm 48 has a symmetry with respect to its plane H1. Each arm 48 is tubular to allow the passage of servitudes through it, it can also be structural or not. Each arm 48 preferably retains its dimensions, such as its cord and its thickness, over the prior art shown in FIGS. 1 and 2. The stationary vanes 44 may be similar to those of the prior art. The annular row of variable pitch vanes 56, 58 comprises first blades 56 which are in the extension of the arms 48 and 20 seconds blades 58 which are located between the first blades 56 and downstream of the blades 44 relative to the blade 44. flow flow. As schematically shown in the drawing, the profiles of the blades 56 are more pronounced than those of the blades 58. The portion of the airflow flowing through the stationary vanes 44 is further rectified than the portion of the flow of air flowing through the vanes. air flowing along the arms 48 due to the curvature of the profile of the blades 44. For the air flow leaving the stator is homogeneous over its entire circumference, it is necessary that the part of the air flow s The trailing edges of the arms 48 are more straightened than that flowing from the trailing edges of the fixed vanes 44. The profiles of the vanes 56, 58 are thus configured so that the portion of the air flow flowing from The trailing edges of the arms 48 are further straightened than that of the trailing edges of the fixed vanes 44, and the airflow leaving the stator is homogeneous over its entire circumference. This air flow advantageously has a velocity triangle 38 substantially identical to that of the prior art (FIG. 2) so that the flow of air seen by the downstream rotor wheel 28 is not affected by the configuration of the stator / input casing according to the invention. Referring now to Figure 5 which shows a more concrete embodiment of a stator 40 according to the invention, which is here an inlet housing 18.
    [0017] As in the aforementioned example, the stator 40 comprises rows of stationary vanes 44, arms 48 and variable-pitch vanes 56, 58, the latter having two distinct profiles. These rows of blades 44, 56, 58 and arms 48 extend between two coaxial annular walls, respectively radially inner 62 and outer 64.
    [0018] Each variable pitch vane 56, 58 comprises at its outer radial end a cylindrical pivot 66 which is mounted in a chimney 68 of the outer wall 64 and which is connected by a link 69 to a control ring (not shown) for the positioning the blade 56, 58 around the axis defined by its pivot 66.
    [0019] Referring now to FIGS. 6 and 7, which represent alternative embodiments of stators 40 ', 40 "according to the invention, which here are intermediate housings 20. As in the above example, each stator 40', 40" comprises rows of stationary vanes 44, arms 48 and variable pitch vanes 56, 58 having two distinct profiles. These rows of blades 44, 56, 58 and arms 48 extend between two coaxial annular walls, respectively radially inner 62 and outer 64. Each variable-pitch blade 56, 58 comprises at its outer radial end a cylindrical pivot 66 which is mounted in a chimney 68 of the outer wall 64 and which is connected by a connecting rod 69 to a control ring (not shown) for the positioning of the blade 56, 58 around the axis defined by its pivot 66.
    [0020] In addition, the outer wall 64 further comprises air discharge means. In the case of FIG. 6, the outer wall 64 comprises, upstream of the arms 48, an annular row of through-slots 70 of discharge air passage and means 72 for closing off these slots 70.
    [0021] In the case of FIG. 7, the outer wall 64 comprises, between the arms 48, at least one through-opening 80 for the discharge air passage and at least one door 82 for closing off this opening 80, the door 82 being pivotally mounted about an articulation axis 84. This hinge pin 84 is located at an upstream end of the door 82 and is for example substantially tangential to a circumference centered on the longitudinal axis of the stator 40 " Preferably, the stator 40, 40 ', 40 "according to the invention is monobloc, that is to say formed in one piece, except of course variable-pitch vanes which must remain mobile. The stator may be made in several steps, for example from a foundry blank which is machined and then mechanically welded. 20
    权利要求:
    Claims (10)
    [0001]
    REVENDICATIONS1. Aircraft turbine engine stator (40, 40 ', 40 ") having an annular array of stationary vanes (44) and an annular array of arms (48), the stator having an axis of revolution (A), characterized in that the trailing edges (42) of the vanes are located substantially in a first plane (P1) transverse to said axis and situated downstream from a second plane (P2) transverse to said axis and passing substantially through the leading edges ( 46) arms.
    [0002]
    2. Stator (40, 40 ', 40 ") according to claim 1, characterized in that the leading edges (50) of the blades (44) are located substantially in a third plane (P3) transverse to said axis and located upstream of the second plane (P2).
    [0003]
    3. Stator (40, 40 ', 40 ") according to claim 1 or 2, characterized in that the median longitudinal planes (H1) of the arms (48) are inclined relative to the longitudinal axis (A) of the stator.
    [0004]
    4. Stator (40, 40 ', 40 ") according to claim 3, characterized in that the median longitudinal planes (H1, H2) of the arms (48) and the blades (44) are substantially parallel.
    [0005]
    5. Stator (40, 40 ', 40 ") according to one of the preceding claims, characterized in that it further comprises an annular row of variable-pitch vanes (56) located directly downstream of the annular row. of arms (48).
    [0006]
    The stator (40, 40 ', 40 ") according to claim 5, characterized in that the annular row of variable pitch vanes (56, 58) comprises first variable pitch vanes (56) which are located substantially in the extension of the arms (48) and second variable pitch vanes (58) which are located between the first variable pitch vanes, the first variable pitch vanes having different airfoils than the second variable pitch vanes 3027053 13
    [0007]
    7. Stator (40, 40 ', 40 ") according to one of the preceding claims, characterized in that it comprises two annular walls respectively inner (62) and outer (64), between which the rows of vanes (44) and arms (48).
    [0008]
    8. Stator (40, 40 ', 40 ") according to claim 7, in dependence on claim 5 or 6, characterized in that each variable-pitch blade (56) comprises at its outer radial end a cylindrical pivot (66). which is mounted in a chimney (68) of the outer wall (64).
    [0009]
    9. Stator (40, 40 ', 40 ") according to claim 7 or 8, characterized in that the outer wall (64) comprises: - upstream of the arms, an annular row of through slots (70) of passage of discharge air and means (72) for sealing these slits, or between arms, at least one through-opening (80) of discharge air passage and at least one shut-off door (82). this opening.
    [0010]
    10. Aircraft turbomachine, such as a turbojet engine or a turboprop engine, characterized in that it comprises at least one stator 20 (40, 40 ', 40 ") according to one of the preceding claims.
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    同族专利:
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    引用文献:
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    ITUA20161507A1|2016-03-09|2017-09-09|Gen Electric|GAS TURBOMOTOR WITH AN AIR BREAKING.|
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    US10385871B2|2017-05-22|2019-08-20|General Electric Company|Method and system for compressor vane leading edge auxiliary vanes|US20180017079A1|2016-07-15|2018-01-18|General Electric Company|Variable-cycle compressor with a splittered rotor|
    FR3082229B1|2018-06-08|2020-07-03|Safran Aircraft Engines|TURBOMACHINE WITH A PARTIAL COMPRESSION VANE|
    FR3082230B1|2018-06-11|2020-08-28|Safran Aircraft Engines|NON-HOODED ROTOR AIRCRAFT ENGINE WITH STATOR VANE ADAPTATION|
    EP3608505B1|2018-08-08|2021-06-23|General Electric Company|Turbine incorporating endwall fences|
    US11149552B2|2019-12-13|2021-10-19|General Electric Company|Shroud for splitter and rotor airfoils of a fan for a gas turbine engine|
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    法律状态:
    2015-10-15| PLFP| Fee payment|Year of fee payment: 2 |
    2016-04-15| PLSC| Search report ready|Effective date: 20160415 |
    2016-10-05| PLFP| Fee payment|Year of fee payment: 3 |
    2017-09-21| PLFP| Fee payment|Year of fee payment: 4 |
    2018-02-09| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170717 |
    2018-09-19| PLFP| Fee payment|Year of fee payment: 5 |
    2019-09-19| PLFP| Fee payment|Year of fee payment: 6 |
    2020-09-17| PLFP| Fee payment|Year of fee payment: 7 |
    2021-09-22| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1459721A|FR3027053B1|2014-10-10|2014-10-10|AIRCRAFT TURBOMACHINE STATOR|
    FR1459721|2014-10-10|FR1459721A| FR3027053B1|2014-10-10|2014-10-10|AIRCRAFT TURBOMACHINE STATOR|
    BR112017006884-2A| BR112017006884A2|2014-10-10|2015-09-29|aircraft turbomachine stator|
    RU2017111463A| RU2706098C2|2014-10-10|2015-09-29|Stator of aircraft gas turbine engine and aircraft gas turbine engine|
    US15/517,096| US10458247B2|2014-10-10|2015-09-29|Stator of an aircraft turbine engine|
    PCT/FR2015/052583| WO2016055715A2|2014-10-10|2015-09-29|Stator of an aircraft turbine engine|
    CN201580053979.1A| CN107109947B|2014-10-10|2015-09-29|The stator of aircraft turbine engine|
    EP15785162.7A| EP3204620B1|2014-10-10|2015-09-29|Stator of an aircraft turbine engine|
    CA2963487A| CA2963487A1|2014-10-10|2015-09-29|Stator of an aircraft turbine engine|
    JP2017518299A| JP6625624B2|2014-10-10|2015-09-29|Aircraft turbine engine stator|
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