Patent Grant
11421537
The present invention concerns the field of turbine engines and, in particular, a turbine engine blade equipped with a cooling circuit intended to cool it.
The technical background comprises patent documents WO-A1-2014/116475, EP-A1-3 214 269 and US-A1-2015/110639.
Turbine engine vanes, in particular high-pressure turbine blades are subjected to very high temperatures which can reduce their lifespan and degrade the performance of the turbine engine. Indeed, the turbine engine turbines are arranged downstream from the combustion chamber of the turbine engine which ejects a hot gaseous flow, which is expanded by the turbines and makes it possible to drive them in rotation to operate the turbine engine. The high-pressure turbine which is placed directly at the outlet of the combustion chamber is subjected to the highest temperatures.
In order to make it possible for the turbine blades to support these severe thermal constraints, it is known to provide a cooling circuit wherein a relatively colder air circulates which is taken from the compressors, the latter being located upstream from the combustion chamber. More specifically, each turbine vane comprises a blade with a pressure side and an suction side which are connected upstream by a leading edge and by a trailing edge.
The cooling circuit comprises several cavities inside the blade between the pressure and suction sides of the blade, some of which communicate together and which are supplied by the cooling air from the base of the blade, some of this cooling air opening into outlet orifices which are placed in the vicinity of the trailing edge. These orifices deliver cooling air jets on the walls of the blade.
It is known that the cooling circuit comprises a partition extending, on the one hand, radially, and on the other hand, transversally in the blade so as to form a first “rising” cavity and a second “descending” cavity arranged successively along the direction of circulation of the cooling air and which communicate together through a curved passage. These cavities and the passage are known by the expression of “trombone” circuit. The curved passage is delimited by the radially external free end of the partition which presents a curvature or turning of the partition. This makes it possible to scan a large surface inside the blade for its cooling.
The partition transversally connects a first wall, exposed to the outside environment of the blade and in particular, to the hot gaseous flows, to a second wall which are opposite. The radially external free end of the partition forms a right angle (90°) with each first and second wall at the level of the intersection of the partition forming a connection zone. The intersection made between the first wall and the partition is highly mechanically constrained. This is due to the presence of high thermal gradients between the outside of the blade and inside of it thus inducing a difference in thermal dilatations between the inside and the outside of the blade and in particular at the level of the connection zone. The lifespan of the vane is impacted and the vibratory capacity of it reduced.
The aim of the present invention is to reduce the mechanical constraints that the blade of the vane is subjected to due to the arrangement of a cooling circuit while avoiding the significant structural modifications of the vane.
This aim according to the invention is achieved thanks to a turbine engine vane comprising a blade extending following a radial axis and a cooling circuit arranged inside the blade, the cooling circuit comprising a first cavity and a second cavity arranged downstream from the first cavity following a direction of circulation of a cooling fluid, the first and second cavities extending radially inside the blade and being separated at least partially by a radial partition having a radially external free end which delimits at least partially a passage connecting the first and second cavities, the radial partition connecting a first wall in contact with the outside environment of the blade to a second wall opposite, substantially following a transversal axis, perpendicular to the radial axis, respectively in a connection zone, at least one connection zone presenting a thickening having a substantially triangular general transversal cross-section.
Thus, this solution makes it possible to achieve the abovementioned aim. In particular, the thickening of the connection zone at the level of one of the walls of the blade being subjected to high thermal and mechanical constraints where the radial partition is connected, makes it possible to reduce the maximum constraint in this place. This configuration makes it possible to reduce constraints by around 15% which is significant and makes it possible to improve, on the one hand, the lifespan of the vane and on the other hand, the vibratory capacities of them. Moreover, such a configuration makes it possible to offset the maximum thermal constraint zone of one of the walls inwards of the blade, which is colder, which further improves the lifespan of the vane. Finally, this solution is simple and economical to implement.
The vane also comprises one or more of the following features, taken individually or in combination:
The invention also concerns a turbine engine turbine comprising at least one turbine engine vane presenting any one of the abovementioned features.
The invention further concerns a turbine engine comprising at least one turbine engine turbine such as mentioned above.
The invention will be better understood, and other aims, details, features and advantages of it will appear more clearly upon reading the following detailed, explanatory description of embodiments of the invention given as purely illustrative and non-limiting examples, in reference to the appended schematic drawings, wherein:
This bypass turbine engine 1 generally comprises a fan 2 mounted upstream from a gas generator 3. In the present invention, and generally, the terms “upstream” and “downstream” are defined with respect to the circulation of gases in the turbine engine and here following the longitudinal axis X (and even left to right in
The gas generator 3 comprises, upstream to downstream, a low-pressure compressor 4a, a high-pressure compressor 4b, a combustion chamber 5, a high-pressure turbine 6a and a low-pressure turbine 6b.
The fan 2, which is surrounded by a fan case 7 carried by a nacelle 8, divides the air which enters into the turbine engine in a primary air flow which passes through the gas generator 3 and in particular, in a primary duct 9, and in a secondary air flow which circulates around the gas generator in a secondary duct 10.
The secondary air flow is ejected by a secondary nozzle 11 at the end of the nacelle as well as the primary air flow being ejected outside of the turbine engine via an ejection nozzle 12 located downstream from the gas generator 3.
The high-pressure turbine 6a, like the low-pressure turbine 6b, comprises one or more stages. Each stage comprises a stator blading mounted upstream from a mobile blading. The stator blading comprises a plurality of stator or fixed vanes, called turbine stator vane, which are distributed circumferentially about the longitudinal axis X. The mobile blading comprises a plurality of mobile vanes which are also distributed circumferentially around a disc centred on the longitudinal axis X. The distributors deviate and accelerate the aerodynamic flow at the outlet of the combustion chamber to the mobile vanes such that these are rotated.
In reference to
The vane 20 comprises a cooling circuit 28 which is arranged inside the blade and which is intended to cool the walls of the blade being subjected to high temperatures of the primary air flow leaving the combustion chamber 5 and passing through it. The cooling circuit 28 comprises several cavities which communicate between them so as to form a trombone type conduit. The latter comprises several curved or turning passages (by around 180°) such that the cooling fluid (here, the cooling air), sweeps all of the vane and from top to bottom following the radial axis. The cooling of the blade is thus optimised.
The base 23 comprises a supply channel 30 which comprises a cooling air inlet 31 taken from downstream from the combustion chamber such that on the low-pressure compressor and which opens into the trombone type conduit. The channel 30 also opens onto a radially internal face 32 of the base of the blade which comprises the cooling air inlet. The cooling circuit also comprises outlet orifices 33 which are arranged in the vicinity of the trailing edge 27 of the blade. The outlet orifices 33 are oriented substantially following the longitudinal axis X and are aligned and distributed regularly substantially following the radial axis. In this manner, the cooling air RF which circulates from the base of the blade passes through the cavities inside the blade and opens into the outlet orifices 33.
As is illustrated in detail in
The first cavity 34 and the second cavity 35 are connected (and/or communicate together) by a first passage 40 of cooling fluid which is located in the upper portion of the radial partition 36, following the radial axis, and which is delimited at least partially by the radially internal free end 37. The closing wall 39 also delimits the first passage 40. The passage is curved. The radially external free end comprises a fillet presenting an external surface 51 and which connects two opposite edges of the radial partition 36 following the chord of the blade.
The radial partition 36 connects a first wall to a second wall opposite the blade substantially following the transversal axis respectively in connection zones 47. In the example represented, the first wall is in contact with the outside environment of the blade being subjected to hot gaseous flows and is formed by the suction side 25. The second wall is itself formed by an internal wall 41 which extends on the one hand, following the radial axis and on the other hand, following a direction substantially parallel to the chord of the vane (or substantially along the longitudinal axis X).
Alternatively, the first wall is formed by the pressure side since it is also subjected to the hot gaseous flows. In this case, the radial partition extends transversally between the pressure side 24 and the internal wall 41 to which it is connected with connection zones respectively. Also, according to another alternative, the partition 36 is connected to the pressure side and to the suction side between which this extends transversally.
Upstream from the first cavity 34, an upstream cavity 43 which extends radially along the vicinity of the leading edge is arranged. A third cavity 45 is located downstream from the second cavity 35. This third cavity is separated from the second cavity by a second partition 46 which extends transversally between the pressure side and a portion of the internal wall 41. The second partition 46 is connected to the closing wall 39 at its external end while its radially internal end is free to form a second cooling air passage. A lower cavity 44 extends following the radial axis and transversally between the internal wall 41 and the pressure side.
The cavities 34, 35, 45 close the trombone type conduit.
In reference to
In very schematic
The thickening 48 presents a predetermined width L measured between the internal surface 25a of the suction side and the edge 49 from which the thickening raises from the radially external end (in particular, from the peak of the peak edge 52). The predetermined width is equal or less than half of the width of the partition following the transversal axis. In the present example, the width of the partition is around 2.40 mm.
The thickening 48 also comprises a first curved surface portion 54 with a first radius R1 which connects the external peripheral surface 50 to the external surface 51 of the radially external free end 37 of the partition. In this example, the first surface portion is concave. The second radius is around 3.6 mm. The thickening 48 also comprises a second curved surface portion 55 with a second radius R2 which connects the external peripheral surface 50 to the internal surface 25a of the suction side. The second radius R2 is greater than the first radius R1. Likewise, the second surface portion is concave. In the present case, the second radius is around 5 mm.
The passage 40 presents a transversal cross-section S which depends on the predetermined width L of the thickening following the transversal axis, on the alpha inclination angle, on the first radius, and on the second radius. This makes it possible to limit the obstruction of the cooling air flow passage from the first cavity to the second cavity. The predetermined cross-section is around 23.7 mm by considering the examples of dimensions described above.
Advantageously, but in a non-limiting manner, the vane is made of a metal alloy and following a production method using the cire perdue casting technique or lost wax moulding. The metal alloy is preferably nickel-based and can be monocrystalline.
This method comprises a first step of producing one or more casting cores. In the present example, the vane comprising a blade provided with several cavities is made from several casting cores forming a casting assembly. The latter comprises, in particular, a first core and a second core which are made of a refractory material such as a ceramic material. The first core presents the complementary shape of the cavities for circulating cooling fluid in the blade. In particular, the first core comprises a joining zone between a first and a second wing extended following a radial height, the joining zone having a complex shape intended to form, as a negative, the connection zone between the first wall of the blade and the partition.
The first and the second cores are assembled together by connecting elements to hold them in position against one another.
Following another step of the method, wax or an equivalent material is injected around cores which are arranged beforehand, advantageously, but in a non-limiting manner, in a press. Once the wax is cooled, a model is obtained comprising cores buried in the wax.
The model is arranged on a column with other similar models so as to form a cluster.
The method further comprises the production of a shell made of a refractory material around the cluster and which acts as a mould. The refractory material is, in the present example, a ceramic. The shell is made by immersing the cluster several times in a ceramic slip.
Following another step of the method, molten metal is poured or cast inside the shell so as to fill the cavities obtained during the removal of the wax in the models in particular, and intended to form the metal parts, here the turbine blades. Indeed, prior to this metal pouring step, a wax removal step is carried out.
When the shell is cooled and solidified, a shakeout step makes it possible to destroy the shell and the cores made of metal parts (vane) so as to make the final vane and the cavities for circulating cooling fluid appear.
In
| Number | Date | Country | Kind |
|---|---|---|---|
| 1903023 | Mar 2019 | FR | national |
| Number | Name | Date | Kind |
|---|---|---|---|
| 5813836 | Starkweather | Sep 1998 | A |
| 8016547 | Propheter-Hinckley | Sep 2011 | B2 |
| 20150110639 | Herzlinger et al. | Apr 2015 | A1 |
| 20210148233 | Dube | May 2021 | A1 |
| Number | Date | Country |
|---|---|---|
| 3214269 | Sep 2017 | EP |
| 2014116475 | Jul 2014 | WO |
| Entry |
|---|
| Preliminary Research Report received for French Application No. 1903023, dated Nov. 19, 2019, 5 pages (1 page of French Translation Cover Sheet and 4 pages of original document). |
| Number | Date | Country | |
|---|---|---|---|
| 20200300095 A1 | Sep 2020 | US |
