• 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 aeronautical turbine blade (10) extending in the radial direction and having a lower surface (22) and an upper surface (24), having a plurality of intrados cavities extending radially from the intrados (22) side of the blade (10), a plurality of extrados cavities extending radially on the extrados (24) side of the blade (10), and at least one central cavity located in the central portion of the blade (10) and surrounded by inlets and extrados cavities, the blade (10) also having a plurality of cooling circuits, in which at least one first cooling circuit (1) comprises: a first cavity (A) and a second cavity (B), the first and the second cavity communicating with one another at an inner radial end (14) and an outer radial end (16) of the blade (10), a third cavity (C) communicating with each other. with the second cavity (B) at the outer radial end (16), a fourth th cavity (D) communicating with the third cavity (C) at the radially inner end (14), the first cavity (A) and the second cavity (B) being configured to be fed jointly with cold air through an opening common air intake at the inner radial end (14), and for the air to flow in the same direction in the radial direction, the first cavity (A) being a cavity intrados , the second cavity (B) being a central cavity, the third cavity (C) and the fourth cavity (D) being extrados cavities.
    公开号:FR3056631A1
    申请号:FR1601421
    申请日:2016-09-29
    公开日:2018-03-30
    发明作者:Adrien Bernard Vincent Rollinger；Romain Pierre Cariou；Thomas Michel Flamme；Sylvain Paquin
    申请人:Safran SA；
    IPC主号:
    专利说明:

    FIELD OF THE INVENTION The present invention relates to the field of blades of aeronautical high pressure gas turbines, More particularly the cooling circuit of these blades, and a gas turbine comprising such blades.
    STATE OF THE PRIOR ART The moving blades of a gas turbine of an aircraft engine, and in particular of the high-pressure turbine, are subjected to the very high temperatures of the combustion gases during engine operation. These temperatures reach values much higher than those which the various parts which are in contact with these gases can withstand without damage, which has the consequence of limiting their service life.
    Furthermore, an increase in the temperature of the gases of the high-pressure turbine makes it possible to improve the efficiency of an engine, therefore the ratio between the thrust of the engine and the weight of an airplane powered by this engine. Consequently, efforts are made in order to produce turbine blades which can withstand higher and higher temperatures, and in order to optimize the cooling of these blades. It is thus known to provide these blades with cooling circuits aimed at reducing the temperature of the latter. Thanks to such circuits, cooling air (or “cold” air), which is generally introduced into the blade by its foot, crosses it by following a path formed by cavities formed in the thickness of the blade before being ejected through orifices opening on the surface of the blade. [0005] Such cooling circuits are said to be “advanced” when they are composed of several independent cavities in the thickness of the blade, or when some of these cavities are dedicated to localized cooling. These cavities make it possible to define a blade compatible with the requirements of engine performance and service life of the parts. However, the cooling circuits draw the "cold" air from the primary stream of the engine, so that the temperature of this cold air, acting as a heat transfer fluid, is lower than the temperature of the air s' flowing on the surface of dawn, called "hot air". This air taken from the primary vein of the engine therefore constitutes a loss and degrades the overall efficiency and performance of the engine.
    Furthermore, the advanced circuits have the disadvantage of causing a large temperature difference between the outer walls of the blade in contact with the vein and the walls at the heart of the blade. These large temperature differences induce stresses that could jeopardize the mechanical strength of the blade in operation and thus minimize its service life.
    [0008] There is therefore a need for an aeronautical gas turbine blade cooling circuit, making it possible to minimize the thermal gradients in order to limit the internal mechanical stresses, and to optimize the cooling of the blade by reducing the flow rate. of air used to cool it, thereby improving the overall efficiency of the turbine.
    PRESENTATION OF THE INVENTION The present disclosure relates to an aeronautical turbine blade extending in the radial direction and having a lower surface and an upper surface, comprising a plurality of lower cavities extending radially on the lower side of the blade, a plurality of upper cavities extending radially from the upper side of the blade, and at least one central cavity located in the central part of the blade and surrounded by lower cavities and upper cavities , the blade also comprising a plurality of cooling circuits, in which at least one first cooling circuit comprises:
    a first cavity and a second cavity, the first and the second cavity communicating with each other at an internal radial end and an external radial end of the blade, a third cavity communicating with the second cavity at the level of the external radial end, a fourth cavity communicating with the third cavity at the internal radial end, the first cavity and the second cavity being configured to be supplied jointly with cold air by a common air intake opening at the internal radial end, and so that the air flows therein in the same direction in the radial direction, the first cavity being a lower cavity, the second cavity being a central cavity, the third cavity and the fourth cavity being upper cavities.
    In the present description, by radial direction, we understand the direction extending from the blade root, that is to say the base of the blade, at the blade head, that is ie the end radially opposite the foot of the blade. Thus, in the present description, the inner radial end designates the blade root, and the outer radial end designates the blade head.
    Furthermore, in the present description, by "upward direction" is meant the direction of air flow in a cavity from the blade root towards the blade head, and by "downward direction", the direction of air flow in a cavity from the blade head to the blade root.
    Each cavity is delimited by a wall. The intrados cavities extend radially, that is to say from the blade root to the blade head, on the intrados side of the blade. In other words, one face of at least part of the wall delimiting each lower surface cavity is in contact with the outside air at dawn on the lower surface side. No side of the wall delimiting each lower surface cavity is in contact with the outside air at dawn on the upper surface side.
    Likewise, the upper cavities extend radially, that is to say from the blade root to the blade head, on the upper surface side of the blade. In other words, a face of at least part of the wall delimiting each upper surface cavity is in contact with the outside air at dawn on the upper surface side. No side of the wall delimiting each upper surface cavity is in contact with the outside air at dawn on the lower surface side.
    The central cavity also extends radially, that is to say from the blade root to the blade head, in the central part of the blade. By central part of the blade, it is understood that no face of the wall delimiting the central cavity is in contact with the air outside at dawn. In other words, the wall delimiting the central cavity is in contact either with one or more cavities of the lower surface and with one or more cavities of the upper surface, with the exception of the walls delimiting the said cavities. '' pressure and extrados.
    A cooling circuit designates a plurality of cavities communicating with each other. In the present description, at least one cooling circuit comprises a first, a second, a third and a fourth cavity.
    The first and the second cavity communicate with each other at the foot of the blade so as to form a first common chamber, and are simultaneously supplied with cold air at the level of this first common chamber. The first and second cavities are then isolated from each other by a wall extending in the radial direction, and communicate again with each other at the level of the blade head so as to form a second common chamber.
    Thus, when the cold air is supplied to the first common chamber, the latter is divided between the first cavity and the second cavity by flowing, in each of these two cavities, in the upward direction.
    The first cavity is a lower surface cavity. Consequently, when the cold air flows into the first cavity, the latter exchanges heat by forced convection with the wall separating the first cavity from the hot air on the lower side. Thus, the air flowing in the upward direction in the first cavity absorbs the heat of the wall while heating up as it approaches the blade head.
    The second cavity is a central cavity. Consequently, the cold air flowing in this cavity does not exchange heat with the hot air outside the blade. Thus, the air flowing in the upward direction in the second cavity, parallel to the air flowing in the first cavity, heats very little as it approaches the blade head. Cold air can thus reach the second common chamber, mixing with the air coming from the first cavity.
    The third cavity communicates with the second cavity at the blade head, and with the fourth cavity at the blade head. The air from the first and second cavities then flows into the third cavity in the downward direction, then into the fourth cavity in the upward direction.
    Consequently, the third and fourth cavities, being upper surface cavities, can be supplied with cold air coming from the second cavity, thus improving the cooling on the upper surface. This configuration also has the advantage of targeting cooling on the desired areas. Indeed, the air moving within the cavities undergoes the Coriolis force linked to the rapid rotation of the dawn. This allows the cold air to be pressed against the outer walls as much as possible, that is to say the walls defining the lower surface and upper surface of the blade, constituting the hottest zones. The heat exchanges are thus optimized at the level of the external walls, this also making it possible to improve the uniformity of the temperatures at the heart of the dawn, to minimize the thermal gradients at the heart of the dawn, and therefore to limit the stresses. internal in the walls forming the cavities, thereby improving the mechanical strength thereof. In addition, the second cavity, acting as a mechanically flexible core, makes it possible to absorb the mechanical stresses generated by the thermal expansion of the external walls in contact with the hot air.
    In certain embodiments, the blade has a plurality of lower surface orifices, each communicating with the first cavity and opening onto the lower surface of the blade.
    These intrados orifices may be holes made in the wall separating the first cavity and the intrados of the blade and distributed in the radial direction over at least part of the first cavity, allowing a certain volume d cold air flowing in the upward direction into the first cavity to be exhausted on the lower surface of the dawn.
    This exhaust air makes it possible to further cool the external face of the lower surface, by generating a cooling film on the lower surface of the blade. In addition, the air circulating in the first cavity gradually heats up in the upward direction, as described above. These intrados orifices allow a large portion of the heated air to be discharged outside the first cavity. Consequently, the air supplying the third cavity, coming from the first and second cavities, consists mainly of cold air coming from the second cavity. This improves the cooling of the upper cavities, that is to say the third and the fourth cavity.
    In certain embodiments, the blade has a plurality of upper orifices, each communicating with the fourth cavity and opening onto the upper surface of the blade.
    These upper orifices may be holes made in the wall separating the fourth cavity and the upper surface of the blade and distributed in the radial direction over at least part of the fourth cavity, allowing a certain volume d the air flowing in the upward direction into the fourth cavity to be evacuated on the upper surface of the dawn.
    This exhaust air makes it possible to further cool the external face of the upper surface wall, by generating a cooling film on the upper surface of the blade. In addition, these upper ports allow air to be drawn along the cooling circuit. Indeed, the pressure on the suction side being much lower than the supply pressure of the cooling circuits, the presence of these orifices makes it possible to naturally create an air flow in the cooling circuit.
    In certain embodiments, the blade comprises at least a second cooling circuit comprising two intrados cavities communicating with one another by a plurality of passages distributed in the radial direction along the blade between these two cavities, one of these two cavities being supplied with cold air by an air intake opening at the internal radial end of the blade.
    The cavity which is supplied with cold air can also communicate with the lower surface of the blade via orifices distributed in the radial direction over at least part of said cavity. Thus, when the cold air flows into this cavity, it exchanges heat by forced convection with the wall separating this cavity from the hot air on the lower surface, and is also evacuated through the orifices, generating a film of cooling on the lower surface of the blade, while also entering the other cavity via the plurality of passages.
    In certain embodiments, the blade comprises at least a third cooling circuit comprising a suction surface cavity and a trailing edge cavity extending radially both on the suction side and on the pressure side of the blade at the trailing edge, the two cavities being supplied with cold air by an air intake opening at the internal radial end of the blade, the upper cavity forming an angle at the level of the 'outer radial end of the blade, so as to extend to the trailing edge of the blade.
    In some embodiments, the cavities of the third cooling circuit communicate with a plurality of trailing edge orifices opening onto the underside face of the blade.
    In some embodiments, the plurality of cooling circuits are independent of each other.
    Independent of each other, it is understood that none of the cavities making up a given cooling circuit communicates with a cavity of another cooling circuit. This allows, for each cooling circuit, to produce targeted cooling on specific areas of the blade, without these circuits interfering with each other.
    In some embodiments, the blade has less than two fine cavities, the fine cavities having a first length greater than or equal to at least seven times a second length in a section perpendicular to the radial direction.
    In certain embodiments, the thickness of each fine cavity is less than or equal to 1.2 mm, the thickness being the distance between two sides of the fine cavity along the first length, in a section perpendicular to the radial direction.
    In some embodiments, each fine cavity extends over at least half of the blade in the radial direction.
    In some embodiments, the blade has at most one fine cavity.
    The presence of a limited number of fine cavities facilitates the process of manufacturing the blades. Indeed, the ceramic cores necessary for the development of the cooling circuits are very fragile due to their geometry linked to the small thickness of the fine cavities. Minimizing the number of these fine cavities overcomes these drawbacks.
    The present disclosure also relates to a gas turbine comprising vanes of the present disclosure.
    BRIEF DESCRIPTION OF THE DRAWINGS The invention and its advantages will be better understood on reading the detailed description given below of different embodiments of the invention given by way of nonlimiting examples. This description refers to the pages of attached figures, on which:
    - Figure 1 shows a perspective view of a turbine blade according to the present invention;
    - Figure 2 is a perspective view symbolically representing the cavities of the different cooling circuits of the blade;
    - Figure 3 is a perspective view symbolically representing the cavities of the first cooling circuit of the blade
    - Figures 4A to 4E show cross sections as shown in Figure 3 for different positions in the radial direction from the blade root to the blade head;
    - Figure 5 shows a cross section of the blade, illustrating the areas where the heat transfers are the most important.
    DETAILED DESCRIPTION OF EXAMPLES OF EMBODIMENT The invention is described below with reference to FIGS. 1 to 5. It is noted that FIGS. 2 and 3 do not represent parts of the blade as such, but represent the dawn cavities. In other words, the lines illustrated in Figures 2 and 3 symbolize the internal walls of the blade delimiting these cavities.
    Figure 1 illustrates a movable blade 10, for example metal, of a high-pressure turbine of a turbomachine. Of course, the present invention can also be applied to other mobile or fixed blades of the turbomachine.
    The blade 10 comprises an aerodynamic surface 12 (or blade) which extends radially between a blade root 14 and a blade head 16. The blade root 14 is adapted to be mounted on a disc of the rotor of the high-pressure turbine, the blade head 16 being radially opposite the blade root 14.
    The aerodynamic surface 12 has four distinct zones: a leading edge 18 disposed opposite the flow of hot gases from the combustion chamber of the turbomachine, a trailing edge 20 opposite the leading edge 18 , a lower surface side face 22 and an upper surface side face 24, these side faces 22, 24 connecting the leading edge 18 to the trailing edge 20.
    At the blade head 16, the aerodynamic surface 12 of the blade is closed by a transverse wall 26. Furthermore, the aerodynamic surface 12 extends radially slightly beyond this transverse wall 26 of so as to form a bowl 28, hereinafter called the dawn bathtub. This bathtub 28 therefore has a bottom formed by the transverse wall 26, an edge formed by the aerodynamic surface 12 and it is open towards the blade head 16.
    According to the example described, the blade 10 has three cooling circuits, independent of each other, for cooling the blade: a first cooling circuit 1, a second cooling circuit 2, and a third cooling system 3.
    The first cooling circuit 1 comprises a first cavity A, a second cavity B, a third cavity C and a fourth cavity D. The first cavity A is a pressure cavity, the second cavity B is a central cavity, the third and fourth cavities C and D are upper cavities.
    The first cooling circuit is supplied with cold air by the cavities A and B at the root of the blade 14. The cold air is air drawn from other engine circuits, colder than air flowing on the side faces of the lower surface 22 and upper surface 24, and acting as heat transfer fluid. The first and second cavities A and B communicate with each other at the level of the blade root 14, in the lower 40% of the blade, preferably 25%, more preferably 10% in the radial direction, so as to form a first common chamber 31 (FIGS. 3 and 4E) extending radially from the blade root 14 over a length L1. The length L1 can represent at most 40% of the total length of the blade. The first and second cavities A and B also communicate with each other at the level of the blade head 16, in the upper 20% of the blade, preferably 15%, more preferably 10% in the radial direction, so that form a second common chamber 32 (FIGS. 3 and 4A) extending radially over a length L2. The length L2 can represent at most 20% of the total length of the blade. Between the common rooms 31 and 32, the cavities A and B are isolated from one another by a wall P extending radially along the blade 10. Consequently, the air coming from the first common room 31 then flows separately and parallel in the upward direction in the cavities A and B (arrows in FIG. 3), up to the second common chamber 32.
    Furthermore, the first cavity A communicates with the pressure face 22 of the blade 10 via a plurality of pressure holes 40, distributed radially along the blade 10. Thus, part of the the air circulating in the first cavity A is evacuated by the orifices 40, so as to create a cooling film on the underside face 22, as well as by a blade head orifice 42 located on the blade head, so as to create a cooling film on the wall 26 of the bathtub 28. The air circulating in the first cavity A which is not evacuated by the orifices 40 or 42 mixes with the air coming from the second cavity B , in the second common room 32.
    In addition, the cavity A can be delimited, in its upper part, by a curved wall A 'extending over 20%, preferably 15%, preferably still 10% of the length of the blade in the radial direction, the curvature of the wall A 'being directed towards the leading edge 18. This curved shape of the wall makes it possible to guide the air flowing in the cavity A towards the following cavities, and to ensure a distribution homogeneous air in the cavities while limiting pressure losses. In addition, the wall P separating the cavities A and B may comprise, in its upper part, a curved part P ', forming an angle with respect to the rest of the wall P, so that this curved part P' is directed towards the leading edge 18. This curved part P 'makes it possible to guide the air flowing in the cavity B towards the cavity C. The curved wall A' and the curved part P 'thus make it possible to facilitate the return of the air coming from cavities A and B towards cavity C, that is to say to facilitate the change of direction of air flow, passing from an upward direction in the cavities A and B to a downward direction in the cavity C. This also makes it possible to limit the pressure losses at the time of this reversal.
    The second and third cavities B and C communicate with each other at the level of the blade head 16, in the upper 20% of the blade, preferably 15%, more preferably 10% in the radial direction, from so as to form a third common chamber 33 (FIG. 3 and 4B) extending radially over a length L3. The second and third chambers 32 and 33 therefore communicate with each other at the level of the blade head 16 (FIGS. 3 and 4A), so that the first cavity A can also communicate with the third cavity C. The air circulating in the third cavity C therefore comes from cavities A and B, and flows in the downward direction.
    The length L3 may preferably be greater than the length L2. Thus, the air flowing in the third cavity C comes mainly from the second cavity B. In addition, the air coming from the first cavity A has been largely evacuated through the lower surface orifices 40 and the orifice blade tip 44. More precisely, at least 75%, preferably at least 80%, more preferably at least 85% of the air flowing in the third cavity C comes from the second cavity B. This has the advantage of retaining cold air within the third cavity C, so as to cool the upper surface 24 of the blade more effectively. In fact, the second cavity B being a central cavity, the air coming from the latter is cooler than the air coming from the first cavity A, the latter being heated by thermal transfers, in particular by forced convection, with the face lower surface 22.
    The third and fourth cavities C and D communicate with each other at the blade root 14 in the lower 10% of the blade, preferably 8%, more preferably 6% in the radial direction, so that form a fourth common chamber 34 (FIGS. 3 and 4E). The air circulating in the fourth cavity D therefore comes from the third cavity C, and flows in the upward direction, that is to say from the blade root 14 to the blade head 16. The fourth cavity D communicates with the upper surface via a plurality of upper surface orifices 44 distributed radially along the blade 10. Thus, part of the air circulating in the fourth cavity D is evacuated by the orifices 44, so creating a cooling film on the upper surface 24, as well as through a blade head orifice 46 located on the blade head 16, so as to create a cooling film on the wall 26 of the bath 28 .
    The first cooling circuit 1 thus extends from the lower surface 22, on the side of the trailing edge 20, to the upper surface 24, on the side of the leading edge 18. This configuration makes it possible to exploit the various effects linked to the rapid rotation of the blade 10, in particular the Coriolis force, to press the air in the places requiring an optimization of the heat transfers, in particular the walls delimiting the faces of intrados or extrados of the interior of dawn. The hatched areas in FIG. 5 indicate the areas where the work of the air is minimized, that is to say where the heat transfers are the least important. The arrows in FIG. 5 indicate on the contrary the orientation of the Coriolis force, in other words the zones where the air is plated and where the heat transfers are optimized. This configuration thus makes it possible to reduce the flow of cold air necessary to cool the blade 10, by targeting the heat transfers on the desired zones.
    The central cavity B thus acts as a mechanically flexible core of the blade. In fact, it makes it possible to compensate for the mechanical deformations in the walls constituting the blade 10 adjacent to the lower surface 22 and upper surface 24, caused by the thermal expansions due to the high temperatures on these faces. This thus makes it possible to limit the external over-stresses on the blade 10.
    The second cooling circuit 2, independent of the first cooling circuit 1, has two lower cavities E and F. The cavity E, adjacent to the cavities A, B, C and D of the first cooling circuit, is supplied in cold air at blade root 14 (Figure 4E). The cavity F is located on the side of the leading edge 18 of the blade 10. The cavities E and F communicate with each other by a plurality of passages 52 distributed in the radial direction along the blade 10 between these two cavities ( Figures 4B and 4D).
    The cavity E communicates with the underside face 22 of the blade 10 via orifices 50 distributed in the radial direction over at least part of this cavity E. Thus, when the cold air flows in this cavity, this exchanges heat by forced convection with the wall separating this cavity from the hot air on the lower surface, and is also evacuated through the orifices 50, generating a cooling film on the lower surface of the blade, all by also entering the other cavity via the plurality of passages 52. The air flowing in the cavity F is evacuated by orifices 54 distributed in the radial direction over at least part of this cavity F.
    The third cooling circuit 3, independent of the first and second cooling circuits 1 and 2, comprises an upper cavity G adjacent to the cavities A, B and C, and a trailing edge cavity H extending radially both on the upper side 24 and on the lower side 22 of the blade on the trailing edge side 20. The cavities G and H are supplied with cold air mutually by an air intake opening at the foot of the dawn 14.
    The upper surface cavity G extends on the one hand radially in a first part of cavity G ', from the blade root 14 to the blade head 16 along the upper surface 24, and extends on the other hand in a direction substantially perpendicular to the radial direction in a second part of cavity G, along the bathtub 28, forming an angle in the direction of the trailing edge 20 (FIG. 2), the second part of cavity G making it possible to cool the transverse wall 26 at the trailing edge 2. In other words, the cavity G extends from the blade root 14 to the trailing edge 20.
    Furthermore, the first part of cavity G 'has a high aspect ratio so that, according to a cross section (Figures 4C and 4D for example), one dimension (length) is at least three greater than another dimension (width), giving it a "slender" or elongated shape. This makes it possible to maximize the exchange surface between the air circulating in the cavity G and the upper surface 24. With the exception of the trailing edge cavity H, the shape of which is determined by the shape of the blade 10 at the level from the trailing edge 20, the upper surface cavity G of the third cooling circuit 3 is the only one, on all of the cavities that the blade 10 comprises, to have such a shape ratio. Limiting the number of cavities having such shape ratios facilitates the dawn manufacturing process.
    The trailing edge cavity H does not extend radially over the entire length of the blade 10, and is limited in length by the second part of the cavity G. Furthermore, the cavities of the third cooling circuit 3 communicate with trailing edge orifices 56 opening onto the pressure face 22 at the trailing edge 20, the trailing edge orifices 56 being distributed radially along the blade 10. These orifices 56 make it possible to evacuate the cold air circulating in these two cavities.
    Although the present invention has been described with reference to specific exemplary embodiments, it is obvious that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the revendications. In particular, the numbers of cooling circuits and cavities making up each of these circuits are not limited to those presented in this example. Therefore, the description and the drawings should be considered in an illustrative rather than restrictive sense.
    It is also obvious that all the characteristics described with reference to a method can be transposed, alone or in combination, to a device, and conversely, all the characteristics described with reference to a device are transposable, alone or in combination, to a method.
    权利要求:
    Claims (8)
    [1" id="c-fr-0001]
    1. Aeronautical turbine blade (10) extending in the radial direction and having a lower surface (22) and an upper surface (24), comprising a plurality of lower cavities extending radially on the lower side (22) of the 'blade (10), a plurality of upper cavities extending radially from the upper side (24) of the blade (10), and at least one central cavity located in the central part of the blade (10) and surrounded by lower and upper cavities, the blade (10) also comprising a plurality of cooling circuits, in which at least a first cooling circuit (1) comprises: a first cavity (A) and a second cavity (B), the first and the second cavity communicating with each other at an internal radial end (14) and an external radial end (16) of the blade (10), a third cavity (C ) communicating with the second cavity (B) at the external radial end (16), a qu third cavity (D) communicating with the third cavity (C) at the internal radial end (14), the first cavity (A) and the second cavity (B) being configured to be supplied jointly with cold air through an opening common air intake at the internal radial end (14), and so that the air flows there in the same direction in the radial direction, the first cavity (A) being a lower cavity , the second cavity (B) being a central cavity, the third cavity (C) and the fourth cavity (D) being upper surface cavities.
    [2" id="c-fr-0002]
    2. Dawn (10) according to claim 1, comprising a plurality of intrados orifices (40), each communicating with the first cavity (A) and opening onto the intrados (22) of the blade (10).
    [3" id="c-fr-0003]
    3. Dawn (10) according to claim 1 or 2, comprising a plurality of upper orifices (44), each communicating with the fourth cavity (D) and opening onto the upper surface (24) of the blade (10 ).
    [4" id="c-fr-0004]
    4. Dawn (10) according to any one of claims 1 to 3, comprising at least a second cooling circuit (2) comprising two intrados cavities (E, F) communicating with each other by a plurality of passages (52) distributed in the radial direction along the blade (10) between these two cavities, one of these two cavities being supplied with cold air by an air intake opening at the internal radial end (14) of dawn (10).
    [5" id="c-fr-0005]
    5. Dawn (10) according to any one of claims 1 to 4, comprising at least a third cooling circuit (3) comprising an upper surface cavity (G) and a trailing edge cavity (H) extending radially both on the upper side (24) and on the lower side (22) of the blade (10) at the trailing edge (20), the two cavities being supplied with cold air by an inlet opening for air at the inner radial end (14) of the blade (10), the upper cavity (G) forming an angle at the outer radial end (16) of the blade (10), so as to extend to the trailing edge (20) of the blade (10).
    [6" id="c-fr-0006]
    6. Dawn (10) according to any one of claims 1 to 5, wherein the plurality of cooling circuits are independent of each other.
    [7" id="c-fr-0007]
    7. Dawn (10) according to any one of claims 1 to 6, comprising less than two fine cavities, the fine cavities having a first length greater than or equal to at least seven times a second length in a section perpendicular to the radial direction .
    [8" id="c-fr-0008]
    8. Gas turbine comprising blades according to any one of the preceding claims.
    1/5
    oo 0 oo 0 oo 0 oo 0 oo 0 oo 0 oo 0 oo 0 oo 00 oo
    F1G.1
    类似技术:
    公开号 | 公开日 | 专利标题
    EP3519679B1|2020-05-20|Turbine blade comprising a cooling channel
    CA2475083C|2011-09-13|Cooling circuits for gas turbine blades
    CA2398659C|2010-05-04|Cooling circuits for gas turbine blade
    CA2398663C|2010-02-23|Improvements to cooling circuits for gas turbine blade
    EP1741875B1|2008-09-17|Cooling circuit for a rotor blade of a turbomachine
    EP3267111B1|2022-02-16|Annular wall of a combustion chamber with improved cooling at the primary and/or dilution holes
    EP0785339B1|1999-05-19|Cooled turbine vane
    FR2678318A1|1992-12-31|COOLED VANE OF TURBINE DISTRIBUTOR.
    FR2692318A1|1993-12-17|Fixed nozzle for distributing hot gases from a turbo-machine.
    FR2571428A1|1986-04-11|HOLLOW BLADES OF TURBINES COOLED BY A FLUID AND ENGINE EQUIPPED WITH SUCH PALES
    FR3021699B1|2019-08-16|OPTIMIZED COOLING TURBINE BLADE AT ITS LEFT EDGE
    EP1452695A1|2004-09-01|Cooled turbine blade having reduced cooling air leakage
    CA2981994A1|2016-09-29|Ceramic core for a multi-cavity turbine blade
    FR3057906A1|2018-04-27|OPTIMIZED COOLING TURBINE TANK
    EP3149281A1|2017-04-05|Turbine blade comprising a central cooling duct and two side cavities connected downstream from the central duct
    EP3610133B1|2021-12-29|Turbine blade having an improved structure
    FR3067388A1|2018-12-14|AUBE WITH IMPROVED COOLING CIRCUIT
    FR3066551A1|2018-11-23|MOVABLE DAWN OF A TURBINE COMPRISING AN INTERNAL COOLING CIRCUIT
    FR2659689A1|1991-09-20|INTERNAL COOLING CIRCUIT OF A TURBINE STEERING BLADE.
    FR3090040A1|2020-06-19|Turbomachine blade with improved cooling
    WO2021181038A1|2021-09-16|Turbomachine hollow blade
    EP3942158A1|2022-01-26|Turbine engine blade provided with an optimised cooling circuit
    FR3095834A1|2020-11-13|Improved cooling turbine engine blade
    FR3067389A1|2018-12-14|TURBINE DAWN WITH AN IMPROVED STRUCTURE
    CA3146412A1|2021-02-04|Turbomachine moving blade with cooling circuit having a double row of discharge slots
    同族专利:
    公开号 | 公开日
    EP3519679B1|2020-05-20|
    EP3519679A1|2019-08-07|
    RU2019112437A|2020-11-02|
    US20190211693A1|2019-07-11|
    BR112019006164A2|2019-06-18|
    RU2019112437A3|2020-11-27|
    JP2019529788A|2019-10-17|
    RU2741357C2|2021-01-25|
    US10844733B2|2020-11-24|
    CA3038615A1|2018-04-05|
    JP6908697B2|2021-07-28|
    FR3056631B1|2018-10-19|
    CN109790754A|2019-05-21|
    WO2018060627A1|2018-04-05|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US20090175733A1|2008-01-09|2009-07-09|Honeywell International, Inc.|Air cooled turbine blades and methods of manufacturing|
    US20140199177A1|2013-01-09|2014-07-17|United Technologies Corporation|Airfoil and method of making|
    EP2899370A1|2014-01-16|2015-07-29|Doosan Heavy Industries & Construction Co., Ltd.|Turbine blade having swirling cooling channel and cooling method thereof|
    WO2015171145A1|2014-05-08|2015-11-12|Siemens Energy, Inc.|Airfoil cooling with internal cavity displacement features|
    WO2015181488A1|2014-05-28|2015-12-03|Snecma|Turbine blade with optimised cooling at the trailing edge of same comprising upstream and downstream ducts and inner side cavities|FR3079869A1|2018-04-05|2019-10-11|Safran Aircraft Engines|HIGH PRESSURE TURBINE BLADE COMPRISING A DEAD CAVITY HAVING A SECTION REDUCTION|
    EP3564486A1|2018-05-02|2019-11-06|United Technologies Corporation|Airfoil having improved cooling scheme|
    FR3094037A1|2019-03-22|2020-09-25|Safran|TURBOMACHINE BLADE EQUIPPED WITH A COOLING CIRCUIT AND LOST WAX MANUFACTURING PROCESS OF SUCH A BLADE|
    WO2020224995A1|2019-05-09|2020-11-12|Safran|Turbine vane of a turbomachine, turbine, turbomachine and associated ceramic core for manufacturing a turbine vane of a turbomachine|SU1287678A2|1984-09-11|1997-02-20|О.С. Чернилевский|Cooled turbine blade|
    US5931638A|1997-08-07|1999-08-03|United Technologies Corporation|Turbomachinery airfoil with optimized heat transfer|
    FR2829175B1|2001-08-28|2003-11-07|Snecma Moteurs|COOLING CIRCUITS FOR GAS TURBINE BLADES|
    US6672836B2|2001-12-11|2004-01-06|United Technologies Corporation|Coolable rotor blade for an industrial gas turbine engine|
    US7008179B2|2003-12-16|2006-03-07|General Electric Co.|Turbine blade frequency tuned pin bank|
    US7131818B2|2004-11-02|2006-11-07|United Technologies Corporation|Airfoil with three-pass serpentine cooling channel and microcircuit|
    US7377746B2|2005-02-21|2008-05-27|General Electric Company|Airfoil cooling circuits and method|
    RU2425982C2|2005-04-14|2011-08-10|Альстом Текнолоджи Лтд|Gas turbine vane|
    US7293961B2|2005-12-05|2007-11-13|General Electric Company|Zigzag cooled turbine airfoil|
    US20080028606A1|2006-07-26|2008-02-07|General Electric Company|Low stress turbins bucket|
    US7581927B2|2006-07-28|2009-09-01|United Technologies Corporation|Serpentine microcircuit cooling with pressure side features|
    US7845907B2|2007-07-23|2010-12-07|United Technologies Corporation|Blade cooling passage for a turbine engine|
    US8297927B1|2008-03-04|2012-10-30|Florida Turbine Technologies, Inc.|Near wall multiple impingement serpentine flow cooled airfoil|
    US9017025B2|2011-04-22|2015-04-28|Siemens Energy, Inc.|Serpentine cooling circuit with T-shaped partitions in a turbine airfoil|
    US9115590B2|2012-09-26|2015-08-25|United Technologies Corporation|Gas turbine engine airfoil cooling circuit|
    US9995148B2|2012-10-04|2018-06-12|General Electric Company|Method and apparatus for cooling gas turbine and rotor blades|
    US8920123B2|2012-12-14|2014-12-30|Siemens Aktiengesellschaft|Turbine blade with integrated serpentine and axial tip cooling circuits|
    EP2754856A1|2013-01-09|2014-07-16|Siemens Aktiengesellschaft|Blade for a turbomachine|
    US8920124B2|2013-02-14|2014-12-30|Siemens Energy, Inc.|Turbine blade with contoured chamfered squealer tip|
    FR3012064B1|2013-10-23|2016-07-29|Snecma|FIBROUS PREFORMS FOR TURBOMACHINE HOLLOW DREAM|
    US20150204237A1|2014-01-17|2015-07-23|General Electric Company|Turbine blade and method for enhancing life of the turbine blade|
    FR3021697B1|2014-05-28|2021-09-17|Snecma|OPTIMIZED COOLING TURBINE BLADE|
    US10040115B2|2014-10-31|2018-08-07|United Technologies Corporation|Additively manufactured casting articles for manufacturing gas turbine engine parts|US10731474B2|2018-03-02|2020-08-04|Raytheon Technologies Corporation|Airfoil with varying wall thickness|
    DE102019125779A1|2019-09-25|2021-03-25|Man Energy Solutions Se|Blade of a turbo machine|
    法律状态:
    2017-04-13| PLFP| Fee payment|Year of fee payment: 2 |
    2018-03-30| PLSC| Publication of the preliminary search report|Effective date: 20180330 |
    2018-08-22| PLFP| Fee payment|Year of fee payment: 3 |
    2019-08-20| PLFP| Fee payment|Year of fee payment: 4 |
    2020-08-19| PLFP| Fee payment|Year of fee payment: 5 |
    2021-08-19| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1601421A|FR3056631B1|2016-09-29|2016-09-29|IMPROVED COOLING CIRCUIT FOR AUBES|
    FR1601421|2016-09-29|FR1601421A| FR3056631B1|2016-09-29|2016-09-29|IMPROVED COOLING CIRCUIT FOR AUBES|
    CN201780059960.7A| CN109790754A|2016-09-29|2017-09-28|Turbo blade including cooling circuit|
    BR112019006164A| BR112019006164A2|2016-09-29|2017-09-28|aeronautical turbine blade, and gas turbine.|
    JP2019517017A| JP6908697B2|2016-09-29|2017-09-28|Turbine blade with cooling circuit|
    CA3038615A| CA3038615A1|2016-09-29|2017-09-28|Turbine blade comprising a cooling circuit|
    US16/336,259| US10844733B2|2016-09-29|2017-09-28|Turbine blade comprising a cooling circuit|
    EP17787225.6A| EP3519679B1|2016-09-29|2017-09-28|Turbine blade comprising a cooling channel|
    PCT/FR2017/052636| WO2018060627A1|2016-09-29|2017-09-28|Turbine blade comprising a cooling circuit|
    RU2019112437A| RU2741357C2|2016-09-29|2017-09-28|Turbine blade comprising cooling system|
    [返回顶部]