Patent Application
20240218802
The invention relates to the stator parts of a turbomachine comprising a blade such as the guide vanes located downstream of a compressor and particularly the fixed-pitch guide vanes.
In an aircraft turbomachine, and particularly the aircrafts intended for the transport of passengers, it is the air propelled by a fan and combustion gases leaving the turbomachine through an exhaust nozzle that exerts a reaction thrust on the turbomachine and, through it, on the aircraft. The circulation of the gases through the turbomachine is influenced by rotating vane assemblies and fixed vane assemblies. The fixed vane assemblies or stator vane assemblies include in particular outlet guide vanes (or OGV), inlet guide vanes (or IGV), and variable pitch vanes (also known as Variable Stator Vane or VSV). Typically, the guide vanes of an aeronautical gas turbine engine each have two (internal and external) platforms which are added onto the vane assembly. These guide vanes form rows of fixed vanes which allow guiding the gas stream passing through the engine according to appropriate speed and angle.
Within a guide vane comprising a plurality of fixed blades, the flow of gases generally takes place between the blades along an upstream-downstream direction. It is known, however, that the area of the blade base can be the site of secondary aerodynamic flows.
For each pair of blades facing each other, a pressure gradient between the pressure face (intrados) of the first blade and the depression face (extrados) of the second blade generates a passage flow (also known under the term crossflow) which transports the gases towards the extrados.
At the end of the blade, that is to say at the junction between the vane assembly and the hub or between the vane assembly and the casing, a corner separation and a corner vortex can occur. This separation generates pressure losses as well as aerodynamic blocking. The latter is problematic in terms of operability. For high incidences of the stream arriving on the guide vane, that is to say when the gas flow direction upstream of the guide vanemakes a significant angle with a direction of the leading edge of the blade, this corner separation increases to the point of causing a detachment of the boundary layer on the blade which can no longer ensure the deflection of the flow.
The reduction in the performances and operability of the compressors is all the greater as the ratio s/c between the circumferential distance separating two blades s and the chord of a blade c is large. For lightweight engines with a reduced number of blades and made more compact axially by shortened chords, this ratio s/c is greater, making the effects all the more problematic.
There is therefore a need for a new geometry for correcting these problems and improving the performances in terms of equipment efficiency, in particular at high incidence of the stream entering the guide vane.
One aim of the invention is to propose a stator part of a turbomachine whose geometry improves the flow of the fluids compared to the prior art.
The aim is achieved within the framework of the present invention thanks to a stator part of a turbomachine comprising a platform, a blade and a fin, the blade and the fin extending from the platform, the platform, an extrados of the blade and the fin defining therebetween a gas flow channel, the channel having a section in a plane normal to an axis of the turbomachine, having a surface area which continuously decreases from upstream to downstream with reference to a general gas flow direction through the turbomachine.
On the one hand, the proposed fin limits the passage flow which is directed towards the extrados. On the other hand, the fin defines between it and the extrados a channel in which the fluid flows. This channel has a section which decreases downstream so that the section seen by the fluid through this channel narrows. By preservation of the flow rate in the channel, the flow of the fluid accelerates downstream in the axial direction. There is therefore an acceleration of the stream on the extrados side, which reduces the thickness of the boundary layer on the extrados side of the blade as well as on the platform. This also reduces the area of low momentum associated with the corner separation responsible for the aerodynamic blocking. This is true over a wide range of incidence, and particularly at high incidences.
Such a stator part is advantageously and optionally supplemented by the following different characteristics taken alone or in combination:
The invention also relates to a turbomachine comprising a stator part as has just been presented and on an aircraft comprising such a turbomachine.
Other characteristics and advantages of the invention will emerge from the following description, which is purely illustrative and not limiting, and should be read in relation to the appended drawings in which:
With reference to
The fan 2 and the low-pressure compressor 3 are driven in rotation by the low-pressure turbine 7 via a first transmission shaft 9, while the high-pressure compressor 4 is driven in rotation by the high-pressure turbine 6 via a second transmission shaft 10.
In operation, a flow of air compressed by the low-pressure and high-pressure compressors 3 and 4 supplies combustion in the combustion chamber 5, whose combustion gas expansion drives the high-pressure and low-pressure turbines 6, 7. The air propelled by the fan 2 and the combustion gases leaving the turbojet engine 1 through an exhaust nozzle downstream of the turbines 6, 7 exert a reaction thrust on the turbojet engine 1 and, through the latter, on a vehicle or machine such than an aircraft (not illustrated).
Downstream of the fan or of a compression stage, the turbomachine can comprise a stage of straightening vanes. Such a stage of straightening vanes can comprise a stator part 20 as presented with reference to
The stator part 20, or the set 20 of stator parts if it is not in one piece, has at least two consecutive blades 24, 26 and a platform 22 from which the blades 24, 26 extend.
The blade 24 has an extrados 25 which faces a pressure face of the blade 26. In operation, the air flows through the stator part in a flowpath defined by the platform 22, the blades 24 and 26 and the wall 23. The flow takes place in the direction of the axis Δ of the turbomachine and from upstream to downstream along the direction of the axis Δ directed towards the reader in
The blades 24 and 26 each have a pressure face and an extrados. The blades 24 and 26 each comprise a leading edge 52, 39 on the upstream side and a trailing edge on the downstream side. The blades define a chord 36 which is the segment connecting the leading edge and the trailing edge. The chord 36 projected on the direction of the axis of the turbomachine defines an axial chord 37.
Each blade has a camber line 41, 43 which is the curve equal to the average between the curve of the extrados and the curve of the pressure face. More specifically, the camber line is formed by all the points located equidistant from the extrados and the pressure face. The distance from a particular point in the extrados (or the pressure face) is defined here as the minimum distance between the particular point and a point in the extrados (or the pressure face).
On each camber line 41, 43, a maximum camber point is defined (reference 35 on the blade 24). At this point, the length of a segment perpendicular to the chord line and connecting a point of the chord line and a point of the camber line is maximum.
The coordinate of the maximum camber point along the axis x is denoted x0 in
There is also defined:
The stator part 20 also comprises a fin 28 which extends from the platform in the same direction and the same direction of extension as the blades 24, 26. The fin is located between the blades 24 and 26. The fin extends over a radial dimension 31 smaller than a height of the blades. In other words, the fin does not extend from the platform 22 to the wall 23 over the entire height of the flowpath separating the platform 22 from the wall 23. The radial dimension 31 of the fin 28 varies between 1% and 40% of this flowpath height. The radial dimension 31 depends on the size of an upstream boundary layer.
The fin 28 extends along the axis Δ of the turbomachine from an upstream end 33 to a downstream end, as illustrated in
The fin 28 has a flank 32 which is located facing the extrados 25 of the blade 24. The intersection of the flank 32 and of a plane normal to the axis Δ of the turbomachine is a ridge 29. This ridge can be straight or curved.
The flank 32 of the fin 28 may have a rectilinear ridge 29 which makes it possible to define an inclination 52 with the platform 22, as represented in
The platform 22, the extrados 25 of the blade 24 and the fin 28 define therebetween a gas flow channel 30. The channel 30 extends from the extrados 25 to the flank 32 of the fin 28 along the circumferential direction θ. The ridge 29 of the flank 32 of the fin 28 is contiguous to the channel 30. The channel 30 extends radially from the platform 22 to the wall 23 over a length equal to the radial dimension 31 of the fin 28.
The channel 30 follows the shapes of the platform 22, the extrados 25 and the flank of the fin 28. The channel 30 does not extend beyond the radial dimension 31 of the fin 28.
The stator part is configured so that the channel 30 has a section, in a plane normal to the axis Δ of the turbomachine, whose surface area decreases continuously from upstream to downstream.
In other words, if two planes normal to the axis Δ of the turbomachine are chosen, the two planes comprising a downstream plane and an upstream plane upstream of the downstream plane, the section of the channel 30 in the upstream plane is always greater than or equal to the section of the channel 30 in the downstream plane.
The continuous decrease in the surface area of the section can be obtained in different embodiments which can possibly be combined with each other.
In a first embodiment, the extrados 25 and the flank 32 of the fin 28 are separated in each normal plane by a distance which decreases from upstream to downstream. In this case, the radial dimension 31 of the constant fin can be kept constant and the shape of the ridge 39 can be kept identical in the different normal planes.
In a second embodiment, the inclination 52 of the ridge 29 relative to the platform 22 decreases from upstream to downstream. The flank of the fin 28 is then oblique and the angle of the flank relative to the platform 22 decreases downstream.
In a third embodiment, the radial dimension 31 of the fin decreases from upstream to downstream. In this case, the distance separating the extrados 25 and the flank 32 can be kept constant and the shape of the ridge 39 can be kept identical in the different normal planes. The second embodiment and the third embodiment can be advantageously combined: the fin decreases in radial dimension downstream and the inclination of the ridge decreases downstream.
Thanks to the reduction in the section of the channel 30 from upstream to downstream, the area where the gas flow presented in the prior art a small momentum is accelerated in the stator part presented here.
Furthermore, as the channel size decreases, the blocking induced by the channel also decreases.
The boundary layer remains attached longer on the extrados 25 of the blade 24, which improves the straightening efficiency of the latter. This effect is significant at high incidence, where the corner separation is usually significant. By limiting the separations and losses of the stator, the flow is better deflected. This makes it possible to limit the deviation between the gas stream and the profile of the straightening vanes at the stator outlet.
The efficiency of the propulsion assembly formed of the rotor and of the stator is improved. This effect is visible even at low incidence, close to the maximum efficiency point for heavily loaded stators—that is to say for stator guide vanes whose ratio s/c is high.
It is thus possible to have a more robust stator, which can increase the operating margin of the compressor.
Optionally to the embodiments previously presented, the upstream end 33 of the fin 28 can be placed in specific areas according to two conditions.
A first condition is that the upstream end 33 can be located axially, that is to say along the direction of the axis Δ of the turbomachine, upstream of the camber point 35 at a distance less than or equal to 30% of the axial chord 37 and downstream of the camber point 35 at a distance less than or equal to 20% of the axial chord 37.
In other words, the upstream end 33 is located between the straight lines of equation x=x1 and x=x2, with the coordinates x1 and x2 introduced previously. The straight lines x=x1 and x=x2 are represented in dotted lines in
In addition to this first condition, the upstream end 33 can be located, according to a second condition, at particular distances from the tangents of the camber lines 41, 43 of the blades 24, 26. More specifically, the tangent T1 to the camber line 41 of the blade 26 is defined at its leading edge 52, and the tangent T2 to the camber line 43 of the blade 24 is defined at its leading ridge 39.
These two tangents T1 and T2 are parallel and a plane simultaneously normal to the two tangents T1, T2 can be defined. According to the second condition, the upstream end 33 is located at a distance from each of the tangents greater than or equal to 5% of the axial chord 37.
The straight lines K1 and K2 define an area therebetween and if the upstream end 33 of the fin 28 is in this area, the second condition is met.
Furthermore, the axial position of the downstream end of the fin can be the axial position of the trailing edge of the blades 24, 26.
The two conditions make it possible to optimize the position of the fin as a function of the maximum curvature area of the blades and to optimize the effect of controlling the separation on the downstream portion of the blade 24, while reducing the disadvantages of the addition of a fin.
| Number | Date | Country | Kind |
|---|---|---|---|
| FR2108792 | Aug 2021 | FR | national |
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/FR2022/051578 | 8/11/2022 | WO |
