FR3082229A1 - TURBOMACHINE WITH A PARTIAL COMPRESSION VANE - Google Patents

Abstract

The invention relates to a double flow turbomachine (1) comprising: - a fan (3) generating an air flow and comprising an annular row of fan blades (13) carried by a rotor (14) of longitudinal axis X , - a separation nozzle (10), downstream of the row of fan blades, separating the air flow into a primary flow circulating in a primary annular vein (6) and into a secondary flow circulating in a secondary vein (7) annular, and - at least one compressor (9) arranged downstream of an inlet of the primary vein. According to the invention, the turbomachine comprises an annular row of movable partial blades (20) extending substantially radially from the rotor and being arranged axially between the annular row or the last of the annular rows of fan blades and the inlet. of the primary vein, each partial blade having a height less than or equal at least to a height of the entry of the primary vein.

Description

TURBOMACHINE WITH A PARTIAL COMPRESSION VANE

1. Field of the invention

The present invention relates to the field of turbomachinery and in particular to double flow in which air flows from upstream to downstream. It relates more precisely to means for compressing the air flow in the turbomachine.

2. State of the art

Double-flow turbomachines are known comprising a movable fan arranged upstream of at least one compressor according to the flow of gases in the turbomachine. The turbomachine comprises a primary stream in which a primary stream circulates and a secondary stream in which a secondary stream circulates, the compressor being installed in the primary stream. These primary and secondary veins are separated by an annular inter-vein casing. The inter-vein casing carries an annular separation spout separating the primary vein from the secondary vein. The blower comprises fan blades each with a free end facing an external casing so as to provide first compression of the incident air flow in the turbomachine which is directed towards the primary stream and on the other hand, to train the air flow which passes through the secondary vein in order to provide a significant component of the thrust. The air flow circulating in the primary stream is conventionally compressed by one or more compressor stages of the turbomachine before entering the combustion chamber. The combustion energy is recovered by one or more turbine stages which participate in driving the compressor stages and the blower.

In order to optimize the performance of the turbomachine, certain turbomachinery are equipped with a low pressure compressor which makes it possible to provide a first compression of the primary flow circulating in the primary stream before passing through a high pressure compressor. However, for turbomachines which only comprise a single compressor, in this case a high pressure compressor, only the fan which is arranged upstream of the gas generator allows the compression of the air flow. The supply of the high pressure compressor is a sensitive aspect of the design of the turbomachine. The problems specific to the latter can translate into constraints for the components upstream of it. Examples of constraints, in particular upstream of the compressor, are the flow rate of the air flow and the distortion of the rotor, in particular at the level of a portion having the shape of a "swan neck". The problem dealt with here is that of the average inlet pressure of the primary flow: in the event that this is insufficient (at the time of conception, or of redesign). One solution is to size the blades of the blower so that they compress the air flow intended for the primary flow more at their feet than at their heads. However, this solution quickly becomes limited, because involving blades that are too loaded at the foot, that is to say that the feet are dimensioned to support aerodynamic forces, for example axial, large, and too lightly loaded at the head. that is to say that the heads are dimensioned to support less aerodynamic forces compared to the feet.

3. Object of the invention

The present invention aims in particular to provide a double-flow turbomachine making it possible to achieve additional compression of the primary air flow while avoiding major modifications to the configuration of the turbomachine.

4. Statement of the invention

This objective is achieved, in accordance with the invention, thanks to a double-flow turbomachine comprising:

- a fan generating an air flow and comprising at least one annular row of fan blades carried by a rotor of longitudinal axis X,

- a separation nozzle, downstream of the row of fan blades, separating the air flow into a primary flow circulating in a primary annular vein and into a secondary flow circulating in a secondary annular vein, and

- At least one compressor arranged downstream of an inlet of the primary stream formed by an annular edge of the separation nozzle, the turbomachine further comprising an annular row of movable partial vanes extending substantially radially from the rotor and being arranged axially between the annular row or the last of the annular rows of fan blades and the inlet of the primary stream, each partial blade having a height less than or equal to a height of the inlet of the primary stream.

By the expression "partial blade", we mean a blade of low height, this height being less than that of a conventional blade, such as the fan blade moving in a vein of the turbomachine, the height being measured between a foot and a blade head along the radial axis.

Thus, this solution achieves the above-mentioned objective. In particular, the annular row of partial vanes makes it possible to precisely supply the primary stream, in which the compressor is arranged and to regulate the supply pressure thereof. The air flow is then subjected to additional compression to supply the primary stream so as to improve the performance of the turbomachine. Such a configuration of the partial vanes also allows a gain in bulk compared to an architecture with a low pressure compressor mounted upstream of the high pressure compressor. In addition, the mass of the turbomachine is very little impacted with the installation of these partial blades. Finally, the arrangement of these partial blades avoids making structural modifications to the fan blades or to the turbomachine in general, which can generate significant costs and lead time.

According to another characteristic of the invention, the height of the inlet of the primary vein is measured, in a direction perpendicular to the radially internal wall, between a radially internal wall of the primary vein and the annular edge of the separation spout.

According to another characteristic of this embodiment, the annular row of partial vanes is arranged in an area of the rotor having a maximum diameter. Such a configuration makes it possible to reduce the size of each partial blade and thus limit the mass of the partial blade stage.

According to yet another characteristic of the invention, each partial blade has a substantially curvilinear free end. Such a configuration does not increase the incidence of the separation spout. The aerodynamic flow is not disturbed.

According to another characteristic of this embodiment, the turbomachine comprises at least one annular row of stator blades situated between an annular row of fan blades of an n-row blower and the annular row of partial blades.

According to another characteristic of the invention, the fan comprises a single annular stage of fan blades.

According to another characteristic of the invention, the fan comprises several annular stages of fan blades arranged along the longitudinal axis X, between each fan blade being arranged an annular row of stator blades, the annular row of blades partial being arranged downstream of the last row of stator vanes.

According to another characteristic of the invention, the turbomachine comprises at least one structural arm extending radially in the primary vein and / or in the secondary vein.

According to another characteristic, the structural arm has a leading edge upstream of a trailing edge, the leading edge being defined in a first plane situated upstream of a second plane in which the annular edge of the separation spout and the annular row of partial vanes being arranged upstream of the structural arm.

According to another embodiment of the invention, the turbomachine comprises a gas generator of longitudinal axis X comprising the compressor and the fan upstream of the compressor, and at least two non-faired propellers driven in rotation by the gas generator via a power transmission system, the propellers being arranged on either side of the longitudinal axis and each propeller having an offset axis relative to that of the gas generator.

The invention also relates to an aircraft comprising a turbomachine having at least any of the above characteristics.

According to a characteristic of this aircraft, the gas generator is arranged at the rear of the fuselage and in the fuselage.

5. Brief description of the figures

The invention will be better understood, and other objects, details, characteristics and advantages thereof will appear more clearly on reading the detailed explanatory description which follows, of embodiments of the invention given by way of purely illustrative and nonlimiting examples, with reference to the appended schematic drawings.

In these drawings:

FIG. 1 is a very schematic and partial view of a double-flow turbomachine to which the invention applies with a stage of partial blades arranged upstream of an inlet of a primary stream of the turbomachine; and

FIG. 2 is a very schematic and partial view of an aircraft equipped with another embodiment of a double-flow turbomachine according to the invention equipped with propellers offset relative to the axis of the gas generator.

6. Description of embodiments of the invention

FIG. 1 illustrates a turbomachine 1 such as an aircraft turbojet engine to which the invention applies. An example of an aircraft to which the invention applies is an airplane. This turbomachine 1 is here a double-flow turbomachine which extends along a longitudinal axis X.

The double-flow turbomachine generally comprises a gas generator 2 upstream of which is mounted a fan 3. In the present invention, and generally, the terms “upstream” and “downstream” are defined in relation to the circulation gases in the turbomachine and with respect to the longitudinal axis X. The terms "external", "exterior", "interior", "internal" and "radial" are defined with respect to a radial axis Z perpendicular to the axis longitudinal X and with regard to the distance relative to the longitudinal axis X. The turbomachine 1 comprises an annular external casing 4 relative to the longitudinal axis X, which surrounds the gas generator 2 and the fan 3.

The turbomachine 1 comprises an air inlet 5 upstream of the fan 3 which generates and compresses a flow of air entering the turbomachine and separating into a primary flow or flow of hot air circulating in a primary annular vein 6 and in a secondary flow or flow of cold air circulating in a secondary annular vein 7. The primary flow and the secondary flow are separated by an annular inter-vein casing 8. The secondary vein extends radially outside the primary vein. The primary flow passes from upstream to downstream a compressor assembly, a combustion chamber and a turbine assembly forming the gas generator while the secondary flow circulates around the gas generator.

In the example illustrated, the turbomachine comprises a high pressure compressor 9, a combustion chamber (not shown) and a high pressure turbine (not shown). The compressor 9 comprises several compressor stages. Here are shown two stages of compressor stator vanes 9a which are alternated along the longitudinal axis with a stage of movable compressor vanes 9b. The stator vanes of the compressor are mounted on the inter-vein casing 8, and in particular a radially internal surface. The movable blades of the compressor are carried by an internal casing 12, here by a radially external wall 23. The high pressure turbine can comprise several turbine stages also for driving the compressor via drive shafts and by removing the combustion gases from the combustion chamber. The combustion gases are expelled into the atmosphere through a nozzle participating in the thrust of the turbomachine.

The primary and secondary veins 6, 7 are coaxial. The secondary stream 7 is delimited radially by the external casing 4 and the annular inter-stream casing 8 surrounding the compressor 9. A separation nozzle 10 makes it possible to separate the primary flow from the secondary flow. The separation spout 10 is arranged upstream of the inter-vein casing and extends the latter upstream, axially. The separation nozzle 10 also includes an annular edge 11 which defines, with the radially external wall 23 of the internal casing 12, the inlet of the primary stream 6. The high pressure compressor 9 is arranged downstream of the air inlet of the primary vein 6. The latter is delimited radially by the annular internal casing 12 and the intervein casing 8.

The fan 3 comprises at least one annular row of fan blades 13 extending radially from a fan rotor 14 of longitudinal axis X. As illustrated in FIG. 1, the fan 3 comprises several stages of fan blades. fan 13 along the longitudinal axis X. In this example, two stages of annular rows of fan blades 13 extend radially from the fan rotor 14. Each fan blade 13 has a leading edge 15, upstream and a trailing edge 16, downstream opposite axially (along the longitudinal axis X). The fan blades 13 each have a foot 17 located in the rotor 14 which is crossed by the fan shaft and a head 18 facing the external casing 4.

The turbomachine also comprises at least one stage of stator vanes 19 or fixed, radial, known as a fan rectifier vane or a fan flow guide vane. The rectifier vane therefore makes it possible to straighten the air flow generated by the blower. In the present invention, we mean by the term "stationary vane" or "stator vane", a vane which is not rotated about the longitudinal axis X of the turbomachine. In other words, this stator vane 19 is separate and contrary to a moving vane or rotor of the turbomachine. Several rectifier vanes 19 are arranged transversely in the air flow. These stator vanes 19 are regularly distributed around the axis X of the turbomachine. Each stator blade 19 extends radially between the fan rotor 14 and the outer casing 4. These are secured to the outer casing 4. In the present example, three annular rows of stator vanes 19 are alternated with the rows annular fan blades 13 along the longitudinal axis X. In particular, a first annular row of stator blades is arranged upstream of the first annular row of fan blades. This first row of stator vanes 19 is arranged immediately at the air inlet 5 of the turbomachine. Likewise, a last annular row of stator vanes 19 is arranged downstream of the last annular row of fan blades 13.

The turbomachine 1 also comprises a row of movable partial blades 20 extending substantially radially from the rotor 14. These partial blades make it possible to provide additional compression of the air flow intended to supply the primary stream to improve the performance of the turbine engine. The annular row of partial vanes 20 is located upstream of the entry of the primary vein. In particular, the row of partial blades 20 is arranged axially between the annular row of fan blades 13 (or the last of the annular rows of fan blades) and the inlet of the primary vein 6. As can be seen in FIG. 1, the annular row of partial vanes 20 is located downstream of the last annular row of stator vanes 19. Each partial vane 20 extends radially from the rotor 14. These each have a first end 21 located in the rotor 14 and a second free end 22 facing the external casing 4. The first and the second ends are radially opposite. The free end 22 of each partial blade 20 is substantially curvilinear so as not to disturb the flow of secondary aerodynamic air.

Advantageously, but not limited to, the annular row of partial vanes 20 is arranged in an area of the rotor 14 having a maximum diameter. According to the example shown in FIG. 1, the partial blades 20 extend radially at the level of the portion of the rotor in the form of a "swan neck". The latter is located upstream of the entry of the primary vein or even upstream of the separation spout. In this way, the dimensions of each partial blade 20 and in particular their height are limited, which limits the impact on the drag. Of course, the annular row of partial vanes can be placed anywhere on the portion of the rotor in the form of a "swan neck" and not necessarily at the level of the maximum diameter thereof. For example, the annular row of partial vanes 20 can be arranged immediately at the entrance to the primary vein, or even radially below the annular edge 11 of the separation nozzle 10.

In particular, the height of each partial blade 20 is less than or equal to the height of the entry of the primary vein. Likewise, the height of each partial blade 20 is less than that of the primary vein 6. All the partial blades 20 have the same height and are also of height less than that of the fan blades. The height of the entry of the primary vein corresponds to the distance measured between the radially external wall 23 of the internal casing 12 and the annular edge 11 of the separation nozzle 10. This distance is measured in particular in a direction which is perpendicular to the radially external wall of the internal casing 12. This direction can be parallel to the radial axis or include a component transverse to the radial axis. In other words, the annular row of partial vanes 20 extends along the radial axis or is slightly inclined relative to the radial axis. In the present example, the radially outer wall 23 extends upstream towards the wall of the rotor 14.

Advantageously, but not limited to, the turbomachine further comprises at least one structural arm 30 extending radially in the primary vein 6 and in the secondary vein 7. In the present example, several structural arms 30 extend around the longitudinal axis X. Each structural arm extends between the outer casing 4 and the rotor 14. Each structural arm 30 has a leading edge 31 and a trailing edge 32 opposite along the longitudinal axis X. The edge of attack 31 of the structural arm 30 is defined in a first plane situated upstream of a second plane in which is defined the annular edge 11 of the separation nozzle 10. Of course, the plane of the leading edge can be arranged downstream of the plane of the annular edge of the separation spout. The annular row of partial vanes 20 is arranged upstream of the row of structural arms 30. The structural arms can make it possible to straighten the flow at the entry of the primary and secondary veins or to pass cables or easements.

FIG. 2 shows another example of a turbomachine to which the invention applies. It is a double-flow turbomachine installed at the rear of a fuselage 35 of an aircraft 40 of axis B. The corresponding reference numbers of the elements of the turbomachine described above are retained in the following description. The turbomachine also comprises at least one gas generator 2 of longitudinal axis X substantially parallel to axis B and which is arranged at the rear of the fuselage 35. This turbomachine 1 is a turbojet engine with non-faired propellers. The partial blades 20 also find their application in these double-flow turbomachines, the gas generator 2 of which is buried in the fuselage to further compress the air flow intended to supply the primary stream. In this embodiment, two turbomachines are mounted on the aircraft 40. Each turbomachine 1 comprises a gas generator 2 of longitudinal axis X and at least one propeller 41 driven by the gas generator 2 via a power transmission system 42. The two gas generators 2 are separate. These are arranged in the fuselage 35, at the rear of the latter, ie at the tail end of the fuselage. More specifically, the gas generators 2 are enveloped by the fuselage 35. The propeller 41 is offset relative to the axis of the gas generator 2 which drives it. In other words, the axis of the propeller 41 is not coaxial with the axis of the gas generator parallel to the axis of the fuselage and is not coaxial with the longitudinal axis either. The two propellers 41 are therefore arranged on either side of the fuselage. Such a configuration makes it possible to reduce the drag of the turbomachine.

Of course, in another embodiment not illustrated, a single gas generator is arranged in the fuselage, at the rear thereof, and which drives two non-faired propellers, with axes offset from the axis of the generator gas, via a power transmission system.

A gas exhaust nozzle 44 is arranged downstream of each gas generator 2 so as to eject the flow of hot gas along the longitudinal axis and at the rear of the fuselage. The exhaust gas nozzle 44 is advantageously located in the aerodynamic continuity of the turbine.

According to the configuration of this embodiment, the air flow entering the turbomachine is compressed in the compressor assembly, then mixed with fuel and burned in the combustion chamber. The combustion gases generated then pass through the turbine assembly to drive, via the power transmission system, the propellers that allow the thrust. The combustion gases are expelled through the nozzle 44 while participating in the thrust of the turbomachine.

The two propellers 41 are not faired. These two propellers are arranged on either side of the fuselage 35 of the aircraft along the longitudinal axis. Here, the turbomachine comprises two pairs of non-keeled and counter-rotating propellers. Each propeller pair also includes an upstream propeller and a downstream propeller. The upstream and downstream propellers of each of the doublets are driven in reverse rotation. As stated above, each pair of propellers is mounted along an axis offset from the axis of the gas generator which drives them. The upstream and downstream propellers of each doublet are arranged in parallel radial planes, which are perpendicular to the longitudinal axis. The rotors of the propeller pairs are each carried by a nacelle 43. The latter is also carried by a pylon 45 extending laterally behind the fuselage 35 with respect to a transverse axis perpendicular to the longitudinal axis. Each propeller 41 has a diameter substantially equal to twice the section of the fuselage, in particular at the rear, where the gas generators are arranged 2. Each nacelle 43 carrying the propeller rotors 41 also has a diameter which is less than the diameter of the fuselage section, at the rear, where the gas generators 2 are arranged. This allows the propellers 41 to be brought closer to the iso performance fuselage. This also makes it possible to facilitate the maintenance of turbomachinery in the event of a breakdown and to reduce the constraints of the pylons. Each propeller pair 41 is driven by a gas generator 2 via a power transmission system 42.

Claims (9)

1. Turbomachine (1) double flow with longitudinal axis X comprising: - a blower (3) generating an air flow and comprising at least one annular row of fan blades (13) carried by a rotor (14) of longitudinal axis X, - a separation nozzle (10), downstream of the row of fan blades (13), separating the air flow into a primary flow circulating in an annular primary stream (6) and into a secondary flow circulating in a secondary annular vein (7), and - at least one compressor (9) arranged downstream of an inlet of the primary stream (6) formed by an annular edge (11) of the separation nozzle (10), characterized in that the turbomachine (1) further comprises an annular row of movable partial blades (20) extending substantially radially from the rotor (14) and being arranged axially between the annular row or the last of the annular rows of fan blades (13) and the inlet of the primary vein (6), each partial blade (20) having a height less than or equal to a height of the inlet of the primary vein (6). 2. Turbomachine (1) according to the preceding claim, characterized in that the annular row of partial blades (20) is arranged in a region of the rotor (14) having a maximum diameter. 3. Turbomachine (1) according to any one of the preceding claims, characterized in that each partial blade (20) has a free end (22) substantially curvilinear. 4. Turbomachine (1) according to any one of the preceding claims, characterized in that it comprises at least one annular row of stator blades (19) located between an annular row of fan blades (13) of a blower (3) with n row and the annular row of partial vanes. 5. Turbomachine (1) according to any one of the preceding claims, characterized in that the fan (3) comprises a single stage of fan blades (13). 6. Turbomachine (1) according to any one of claims 1 to 4, characterized in that the fan (3) comprises several stages of fan blades (13) arranged along the longitudinal axis X, between each fan blade (13) being arranged an annular row of stator vanes (19), the annular row of partial vanes (20) being arranged downstream of the last row of stator vanes (19). 7. Turbomachine (1) according to any one of the preceding claims, characterized in that it comprises at least one structural arm (30) extending radially in the primary stream (6) and / or in the secondary stream (7 ). 8. Turbomachine (1) according to any one of claims 1 to 7, characterized in that it comprises a gas generator (2) of longitudinal axis X comprising the compressor (9) and the blower (3) upstream of the compressor (9), and at least two non-fairing propellers (41) rotated by the gas generator (2) via a power transmission system (42), the propellers (41) being arranged on both sides other of the longitudinal axis X and each propeller (41) having an offset axis relative to that of the gas generator (2). 9. Aircraft (40) characterized in that it comprises at least one turbomachine (1) according to any one of the preceding claims.