Patent Application
20220205366
The present invention relates to the field of turbomachines and more particularly to a deicing system for a separation nozzle of an aeronautical turbomachine.
In an aeronautical turbine of the two spool and dual flow type, the flow streams of the primary flow and of the secondary flow area separated downstream of the fan by a separation nozzle. Within the primary stream, at the inlet of the lowpressure compressor (also commonly called a “booster”), are located a set of fixed inlet guide vanes (also called IGV). In certain phases of flight and on the ground, icing atmospheric conditions can be encountered by the turbomachine, particularly when the ambient temperature is sufficiently low and in the presence of high humidity. Under these conditions, ice can be formed on the separation nozzle and on the inlet guide vanes. When this phenomenon occurs, it can lead to the partial or total obstruction of the primary stream, and to the ingestion of detached blocks of ice into the primary stream. An obstruction of the primary stream causes insufficient feeding of the combustion chamber which can then be extinguished out or prevent the acceleration of the engine. In the event of the detachment of blocks of ice, the latter can damage the compressor located downstream and also lead to the extinction of the combustion chamber. To avoid the formation of ice on the separation nozzle, techniques are known consisting of extracting hot air in the primary stream at a compressor and injecting it inside the separation nozzle. The hot air injected into the separation nozzle can then be routed inside the nozzle to bores or grooves configured to inject the hot air into the primary stream, which can also deice the inlet guide vanes. The necessary flow rate of hot air for deicing the separation nozzle is high. This extraction of hot air can reduce the performance and operability of the turbomachine.
It has seemed desirable to be able to increase the effectiveness of the deicing of the nozzle.
One known solution consists of reducing the volume inside the nozzle, and thus reduce the heat losses inside the nozzle. It is thus known to add an annular baffle in the cavity of the nozzle. The baffle allows reducing the volume of the cavity of the nozzle and orienting the hot air toward the zones of interest for deicing. However, the addition of a baffle (and of its different attachment elements) makes the nozzle heavier, which is manifested by an increase of the fuel consumption of the turbine during operation. It would therefore be desirable to be able to increase the effectiveness of the deicing of the separation nozzle without however increasing the extraction of hot air in a pressurized portion of the turbomachine, without increasing the mass of the nozzle.
According to a first aspect, the invention relates to a separation nozzle between a primary flow and a secondary flow of a dual flow turbomachine. The nozzle has a single-piece structure and comprises an outer annular wall, an inner annular wall, a radial annular wall and an inner annular baffle, defining a first cavity between the outer annular wall and the inner annular baffle, and a second cavity between the inner annular wall, the radial annular wall and the inner annular baffle.
In a particularly advantageous manner, the deflector allows reducing the inner volume of the nozzle in which the hot air circulates. This arrangement therefore allows reducing the heat losses and thus reducing the extraction of hot air. In addition, the baffle allows guiding the hot air within the nozzle.
Moreover, the single-piece structure allows dispensing with numerous connecting parts and therefore reducing the mass of the nozzle compared to known devices. In addition, the mechanically consistent assembly which the single-piece structure constitutes can allow refining the assembly of the walls of the nozzle and further reducing its mass.
Thus, the invention allows increasing the effectiveness of the deicing of the separation nozzle without however increasing the extraction of hot air in the pressurized portion of the turbomachine, without increasing the mass of the nozzle.
The outer annular wall can have at a junction region with the inner annular wall a series of radial holes.
The nozzle can have at least one axial rib between the inner annular wall and the inner annular baffle.
According to one particular arrangement, the nozzle can have a plurality of axial ribs, each coplanar with an axis of revolution of the nozzle.
The beak can have at least one radial rib between the radial annular wall and the inner annular baffle.
According to one particular arrangement, the nozzle can have a plurality of radial ribs, each coplanar with an axis of revolution of the nozzle.
The nozzle can have at least one air cell formed at least partially in the radial annular wall.
The radial annular wall can have a bore leading into the at least one air cell.
The radial annular wall can have at least one oblong opening adapted to accommodate an injector leading into the second cavity.
According to a second aspect, the invention relates to a straightener for an aeronautical turbomachine, which has a single-piece structure formed by additive manufacture, comprising a nozzle having: (i) the single-piece structure comprising an outer annular wall, an inner annular wall, a radial annular wall and an inner annular baffle, (ii) the first cavity between the outer annular wall and the inner annular baffle, (iii) the second cavity between the inner annular wall, the radial annular wall and the inner annular baffle. According to a third aspect, the invention relates to a method for manufacturing a straightener of an aeronautical turbomachine having a single-piece structure formed by additive manufacturing and comprising a nozzle having: (i) a single-piece structure comprising an outer annular wall, an inner annular wall, a radial annular wall and an inner annular baffle, (ii) a first cavity between the outer annular wall and the inner annular baffle, (iii) a second cavity between the inner annular wall, the radial annular wall and the inner annular baffle.
The method can comprise a step of manufacturing the nozzle beginning with the radial annular wall.
Other features and advantages will still be revealed by the description that follows, which is purely illustrative and not limiting, and must be read with reference to the appended figures in which:
With reference to
According to the embodiment presented here, the nozzle 1 is an integral part of a straightener 10 of the primary flow. The nozzle 1 and the straightener 10 are axially symmetrical parts. It is thus understood that the nozzle 1 forms a substantially cylindrical element inside which passes the primary flow, and outside (around) which passes the secondary flow. For the continuation of the description, an axis of revolution X of the straightener 10 (and of the nozzle 1) is defined, and a radial axis Z, substantially perpendicular to the axis of revolution X, shown in
According to a radial direction Z progressing from the interior (closest to the axis of revolution X) toward the exterior (farthest from the axis of revolution X), the straightener 10 comprises successively: an inner ferrule 101, vanes 102 and the nozzle 1.
In a particularly advantageous manner, the nozzle 1 also has a single-piece structure. As described hereafter, the nozzle 1 is preferably formed by additive manufacturing. The nozzle 1 comprises an outer annular wall 12, an inner annular wall 13, a radial annular wall 14 and an inner annular baffle 16. When passing through the nozzle 1 in said radial direction Z, the inner wall 13, the inner annular baffle 16 and the outer annular wall 12 are encountered in succession. A section of the nozzle 1 in a plane XoZ (as can be seen in
The inner annular wall 13 and the outer annular wall 12 join moving upstream (i.e. toward the fan) to form the “nozzle” in functional terms. A junction region of the outer annular wall 12 and the inner annular wall 13 is defined.
The outer annular wall 12 is preferably slightly curvilinear, particularly domed (convex), so as to improve the overall aerodynamics of the nozzle 1.
Between the outer annular wall 12 and the inner annular deflector 16, the nozzle 1 has a first cavity 17.
Between the inner annular wall 13, the radial annular wall 14 and the inner annular baffle 16, the nozzle 1 has a second cavity 18.
In other words, the nozzle 1 is substantially divided into two by the annular inner baffle 16, this defining the two cavities 17, 18. It is understood in fact that the nozzle 1 is substantially hollow (with the exception of a zone in proximity to the radial annular wall 14, see below).
To this end, the inner annular deflector 16 extends from the junction region of the outer annular wall 12 and of the inner annular wall 13 to a junction region of the outer annular wall 12 and the radial annular wall 14. It preferably has an angled shape so that the first cavity 17 occupies the major portion of the volume of the nozzle 1, the second cavity 18 following essentially the radial annular wall 14, then the inner annular wall 13. The second cavity has a first portion 18a between the inner annular wall 13 and the inner annular baffle 16, and a second portion 18b between the radial annular wall 14 and the inner annular baffle 16. It is specified that the two portions 18a and 18b of the second cavity 18 communicated with one another and define a single volume.
With reference in particular to
In addition, preferably, the nozzle 1 comprises a series of axial ribs 22 between the inner annular wall 13 and the inner annular baffle 16, extending in the first portion 18a of the second cavity 18. It is specified that each of the axial ribs 22 is coplanar with the axis of revolution X, i.e. in the plane XoZ.
Likewise, the nozzle 1 comprises a series of radial ribs 24 extending between the radial annular wall 14 and the inner annular baffle 16, extending in the second portion 18b of the second cavity 18. It is specified that each of the radial ribs 24 is coplanar with the axis of revolution X, i.e. again in the plane XoZ.
What is meant here by “axial” and “radial” is simply their main extension direction. Moreover, each axial rib 22 can be coplanar with a radial rib 24. It is understood that the axial and radial ribs 22, 24 define azimuthal partitioning (i.e. sectors) of the second cavity 18, but incomplete ones (i.e. the ribs 22 and 24 nevertheless remains spaced and advantageously do not touch one another), so that at a junction region of the inner annular wall 13 and the radial annular wall 14 (i.e. at the junction of the first and second portions of the second cavity . . . ) the second cavity 18 is not ribbed, allowing an azimuthal communication. Similarly, the axial ribs 22 do not extend until the end of the second cavity, so as to also allow azimuthal communication at the level of the holes 20. In a particularly advantageous manner, the axial 22 and radial 24 ribs have a dual function of mechanical reinforcement and guiding the flow of hot air.
In fact, the axial 22 and radial 24 ribs allow stiffening the nozzle 1, which allows avoiding a possible collapse of the nozzle 1. The axial 22 and radial 24 ribs advantageously allow optimizing the mass of the nozzle 1 by allowing refining the thickness of the inner annular baffle 16, of the radial annular wall 14 and of the inner annular wall 13. It is understood that this mass optimization relies on a compromised between the addition of mass of the ribs and the reduction of thickness of the walls and of the baffle that they allow. Moreover, during the manufacture of the nozzle 1, according to an additive manufacturing method, the axis 22 and radial 24 ribs allow guaranteeing the good mechanical strength of the nozzle 1 during manufacture,
As will be detailed, in operation, the axial 22 and radial 24 ribs allow guiding the flows of hot air to deice the nozzle 1.
Moreover, as can be observed in particular in
Thus, it is remarkable that the formation of the nozzle 1 by additive manufacturing allows obtaining a single-piece structure, but also allows optimizing the geometry of the nozzle 1 to have a better ratio between mass and resistance. In this particular case, the air cells 28a, 28b, 28c would be very difficult to form other than by using additive manufacturing.
The radial annular wall 14 can have bores 30 leading into the first and second air cells 28a and 28b. The bores 30 advantageously allow evacuating a portion of the powder resulting from the additive manufacturing of the nozzle 1.
As shown in
Moreover, the radial wall 14 can have a plurality of attachment bores 35.
Manufacturing Method
In a particularly advantageous manner, the straightener 10 is manufactured by means of an additive manufacturing method.
Thus, the straightener 10 is manufactured by successive additions of melted powder, layer by layer. As previously disclosed, this manufacturing method allows obtaining a single-piece part having a specific geometry.
Preferably, the straightener 10 is manufactured beginning with the radial annular wall 14 of the nozzle, in a progression direction (i.e. of addition of layers of material) substantially parallel to the axis of revolution X.
Operation
An injector (not shown) can be connected to each oblong opening 33. The injectors can blow hot air into the second cavity 18.
In a particularly advantageous manner, the inner annular deflector 16 allows reducing the inner volume of the nozzle 1 by dividing it into two cavities. Thus, the volume in which the hot air circulates is reduced, which reduces heat loss in the nozzle 1 and allow reducing the extraction of hot air. In addition, the inner annular baffle 16 allows orienting the hot air toward the zones of interest for deicing.
The heat radiation of the hot air inside the nozzle 1 allows deicing the nozzle 1. The hot air circulating in the second cavity 18 is then distributed by the radial holes 20 to join the primary stream and deice the vanes 102.
Thus, the invention allows effectively deicing the nozzle without however increasing the extraction of hot air in a pressurized portion of the turbomachine and without increasing the mass of the nozzle.
| Number | Date | Country | Kind |
|---|---|---|---|
| 1904065 | Apr 2019 | FR | national |
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/EP2020/060453 | 4/14/2020 | WO | 00 |
