Transonic compressors for turbines

Abstract

The present disclosure relates to a transonic compressor (24) of a turbine (2), comprising exactly two or exactly three annular rows of rotor blades (40, 44) and exactly one or exactly two annular rows of intermediate stator blades (41) between the respective two rows of rotor blades (40, 44), wherein the flow velocity of the rotor (44) downstream of the intermediate stator blades (41) is less than or equal to Mach number 0.9 over a radial portion of the blade (44) extending over the radially inner 40% of the blade (44), less than or equal to Mach number 1 over a radial portion extending over the radially inner 80% of the blade (44), and less than or equal to Mach number 1.05 over a radial portion extending over the radially outer 20% of the blade (44).

Description

Technical Field

[0001] The present invention relates to the design of turbomachinery compressors and, in particular, to the geometry of stator and rotor blades of transonic compressors. Background Art

[0002] EP 2 669 475 A1 describes an axial flow turbine and focuses in particular on S-shaped geometries for subsonic blades, ie compressor rotor speeds that are particularly suitable for blade speeds below the speed of sound at any point on the blade.

[0003] To reduce the weight and size of a turbine, the compressor, especially the low-pressure compressor, can be made more compact by reducing the number of compression stages that make up the compressor. However, to achieve the same compression ratio with fewer stages, the rotor must rotate faster.

[0004] When the speed of the rotor blades exceeds the speed of sound, the geometry disclosed in EP2669475A1 is no longer suitable. In particular, the relative speed of the air flow at the rotor inlet becomes higher than the speed of sound, and flow blocking and aerodynamic losses caused by shock occur.

[0005] These phenomena also have a feedback effect towards the upstream direction: the blades of the downstream rotor are positioned radially closer to the axis of rotation and will generate a blockage at the blade heads, which will disturb the operation of the blade bases of the upstream rotor in feedback.

[0006] Therefore, existing compressors are not suitable for transonic speeds.

[0007] Thus, the advantage of reducing the number of compression stages in such compressors is mitigated by the range of available speeds being not as high as required to obtain good compressor efficiency. Summary of the Invention

[0008] Technical issues

[0009] The technical problem underlying the present invention is to provide a compressor design which allows a wider range of rotor speeds while avoiding the risk of flow blockage at high speeds.

[0010] Technical Solutions

[0011] The invention relates to an aircraft turbomachine transonic compressor comprising exactly two or exactly three annular rows of rotor blades and respectively exactly one or exactly two intermediate annular rows of stator blades between two consecutive rows of rotor blades, wherein the angle of the air flow at the outlet of an intermediate stator blade of a single annular row of stator blades or of at least one of the two annular rows of stator blades, in particular the most upstream one, is on average greater than 15° over the height of these blades and greater than 15°, preferably greater than 20°, over a radial portion of these stator blades comprised between 60% and 80% of the height of these blades.

[0012] "Transonic" refers to a compressor in which the rotor blades have a speed approaching Mach number 1.2 to 1.4 (at the compressor inlet) and 1.4 to 1.7 within the compressor.

[0013] Thus, by diverting the air flow at the outlet of the intermediate stator blades, the relative velocity at the rotor inlet downstream of the intermediate stator blades can be controlled, thus avoiding choking phenomena.

[0014] The value of 20° is critical because below this value the relative velocity of the fluid becomes too high and clogging is likely to occur.

[0015] The invention can also be defined by the inlet and outlet angles of the intermediate stator blades: a transonic compressor for an aircraft turbine comprising exactly two or exactly three annular rows of rotor blades and exactly one or exactly two annular rows of intermediate stator blades, respectively, between two consecutive rows of rotor blades, a compressor in which the outlet angle of the intermediate stator blades of a single annular row of stator blades or of at least one of the two annular rows of stator blades, in particular the most upstream one, is greater than 9° on average over the height (H) of these blades and greater than 9°, preferably greater than 14°, over the radial portion of these stator blades comprised between 60% and 80% of the height of these blades.

[0016] As described below, under nominal conditions (blade tip speed between 360 m / s and 420 m / s), the difference between the blade exit angle and the flow direction may be of the order of 5° or 6°.

[0017] According to an advantageous embodiment of the invention, the compressor comprises an annular row of downstream stator blades, downstream of the most downstream annular row of rotor blades, the difference between the angle of the inlet air flow of these blades and the angle of the outlet air flow of these blades being less than 40° over the radial portion extending from 10% to 90% of their radial height and less than 50° over the radial portion below 10% and above 90% of their radial height.

[0018] This limited curvature of the most downstream stator of the compressor allows a well-controlled flow within the structural casing directly downstream of the compressor, and therefore also to the high-pressure compressor, and avoids feedback blockages at the last rotor.

[0019] According to an advantageous embodiment of the invention, the two or three rows of rotor blades are two rows of rotor blades comprising one upstream row of blades and one downstream row of blades, the average value of the difference between the angles of the inlet and outlet flow along the entire height of the upstream blades being greater than or equal to 1.4 times, preferably greater than or equal to 1.5 times, the average value of the difference between the angles of the inlet and outlet flow along the entire height of the downstream blades.

[0020] The arithmetic mean of the angular differences along the entire height of the blade gives a picture of the aerodynamic load the blade is subject to. The greater the difference, the more the flow is deflected. This translates to cambered blades, and for a given incoming airflow, the more the flow is deflected, the greater the load on the blade will be.

[0021] Therefore, the first upstream rotor will be the most aerodynamically loaded. This allows the subsequent rotors to have less camber and less aerodynamic load, thus reducing the risk of subsequent rotors blocking while maintaining the nominal compression ratio of the entire compressor. The risk of blocking on the first rotor is less of an issue because no rotor further upstream would be prevented from operating properly at its full height.

[0022] According to an advantageous embodiment of the invention, the two or three rows of rotor blades are three rows of rotor blades, including one row of upstream blades and one row of intermediate blades, the average value of the difference between the inlet and outlet flow angles along the entire height of the upstream blades being greater than or equal to 1.2 times the average value of the difference between the inlet and outlet flow angles along the entire height of the intermediate blades. The three rows further include one row of downstream blades, and optionally, the average value of the difference between the inlet and outlet flow angles along the entire height of the upstream blades being greater than or equal to 1.4 times the average value of the difference between the inlet and outlet flow angles along the entire height of the downstream blades. Thus, in the same manner as in the two-row rotor compressor described above, the downstream rotor blades are less stressed and less likely to generate aerodynamic blockage.

[0023] According to an advantageous embodiment of the invention, the velocity of the flow relative to the rotor encountered at the outlet of the middle stator row is less than or equal to Mach number 0.9 over the radial portion of the rotor blades extending radially inside 40% of the radial height of the blades, less than or equal to Mach number 1 over the radial portion extending radially inside 80% of the radial height of these blades and less than or equal to Mach number 1.05 over the radial portion extending radially outside 20% of the radial height of these blades.

[0024] According to an advantageous embodiment of the invention, the flow direction at the outlet of the rotor row is inclined by more than 40° relative to the axis of rotation of the blade, preferably by an average of 45° to 50° over the height of the blade.

[0025] According to an advantageous embodiment of the invention, the exit velocity of the most upstream rotor blade is less than Mach number 0.8.

[0026] According to an advantageous embodiment of the invention, the two or three rows of rotor blades comprise an upstreammost row and a downstreammost row, and the enthalpy change between the outlet and inlet of the upstreammost row of blades is at least 1.4 greater than the enthalpy change between the outlet and inlet of the downstreammost row of blades.

[0027] The invention also relates to a turbomachine comprising a low-pressure compressor, a high-pressure compressor, a structural casing upstream of the low-pressure compressor and an intermediate structural casing between the low-pressure compressor and the high-pressure compressor, the structural casings delimiting a gooseneck-shaped air flow, wherein the low-pressure compressor conforms to one of the above-described embodiments and is arranged directly downstream of the upstream casing and directly upstream of the intermediate casing.

[0028] The turbomachine may include a fan and a gearbox disposed between any compressor and the fan.

[0029] The term "direct" as used herein means that there are no annular rows of blades (rotor blades or stator blades) between the casing and the compressor.

[0030] The invention also relates to the use of a compressor or turbine as described above, wherein the rotor blades are rotated at a speed such that their radially outer tips move at a speed between 360 m / s and 420 m / s.

[0031] According to an advantageous embodiment of the invention, the speed of the rotor blade is greater than 340 m / s on a radially outer portion extending over at least 30% of the height of the rotor blade.

[0032] According to an advantageous embodiment of the invention, the flow direction directly downstream of the intermediate stator blade is inclined relative to the axis of rotation of the rotor blade by an average angle of at least 15° along the blade height and by an average angle of at least 20° along the radial portion of the stator blade between 60% and 80% of the blade height.

[0033] According to an advantageous embodiment of the invention, the speed of the flow relative to the rotor it encounters at the outlet of the middle stator row is less than or equal to Mach number 0.9 over the radial portion of the blade extending radially inside 40% of the radial height of the blade, less than or equal to Mach number 1 over the radial portion extending radially inside 80% of the radial height of the blade, and less than or equal to Mach number 1.05 over the radial portion extending radially outside 20% of the radial height of the blade.

[0034] According to an advantageous embodiment of the invention, the two or three rows of rotor blades comprise an upstream-most row and a downstream-most row, and the load on the upstream-most row is at least 1.4 greater than the load applied to the downstream-most row.

[0035] "Load" is the increase in enthalpy divided by the square of the rotational speed. The increase in enthalpy is measured as the temperature increase between the blade inlet and outlet. As described below, this increase is related to the deviation in air flow.

[0036] Advantages of the present invention

[0037] The present invention is particularly advantageous in that it allows the compressor to be operated with only two or three high speed compression stages due to the control of relative speed near the rotor downstream of the compressor's intermediate stator. BRIEF DESCRIPTION OF THE DRAWINGS

[0038] Figure 1 The turbine is schematically depicted;

[0039] Figure 2 The rotor / stator stage with speed diagram is shown;

[0040] Figure 3 The rotor blades are shown in an isometric view;

[0041] Figure 4 An isometric view of a stator blade is shown;

[0042] Figure 5 An example of a two-stage compressor is shown;

[0043] Figure 6 An example of a three-stage compressor is shown. DETAILED DESCRIPTION

[0044] In the following description, the terms "inner" and "outer" refer to positions relative to the turbine's axis of rotation. The axial direction is along the turbine's axis of rotation. The radial direction is perpendicular to the axis of rotation. Upstream and downstream refer to the prevailing direction of flow in the turbine.

[0045] The accompanying drawings are not drawn to scale. In particular, the thickness is exaggerated to facilitate reading of the drawings.

[0046] The term "height" refers to the dimension measured along the direction of the largest dimension of the blade, which may be radial or slightly inclined to the radial direction due to the inclination of the air flow.

[0047] The chord is the line segment joining the leading and trailing edges in a plane perpendicular to the radius. The camber is the mid-curve joining the leading and trailing edges at equal distances from the pressure and suction sides.

[0048] Figure 1An axial-flow turbomachine is shown in a simplified form. Here, it is a twin-flow turbojet engine 2.

[0049] The turbine 2 has an axis of rotation 4 . Due to a circular separating nozzle 12 , the annular inlet 6 is divided into a primary duct 8 and a secondary duct 10 .

[0050] The primary flow 14 and the secondary flow 16 pass through these ducts 8 and 10 respectively and merge at the outlet of the turbine 2. The primary flow 14 and the secondary flow 16 are coaxial annular flows and cooperate with each other. They are guided by the inner and outer walls of the turbine 2.

[0051] The secondary flow 16 is accelerated by a fan 18 disposed at the inlet 6 to generate thrust for the aircraft. Straightening vanes 20 may be disposed in the secondary flow duct 10 and configured to increase the axial component of the secondary flow velocity vector. The fan 18 is disposed upstream of the primary duct 8 and the secondary duct 10.

[0052] The turbomachine 2 comprises a compression zone formed by two compressors 24, 26, a combustion chamber 22 and expansion zones 28, 32. The compressors 24, 26 are composed of a low-pressure compressor 24 and a high-pressure compressor 26.

[0053] A high-pressure compressor 26 is located at the inlet of the combustion chamber 22 .

[0054] Downstream of the combustion chamber 22, the turbomachine 2 may have a high-pressure turbine 28 coupled to a high-pressure shaft 30, followed by a low-pressure turbine 32 coupled to a low-pressure shaft 34. The latter may rotate independently of the high-pressure shaft 30. These turbines 28, 32 form an expansion zone for the primary flow 14.

[0055] During operation, the mechanical power received by the turbines 28, 32 is transmitted to the shafts 30, 34, which sets into motion the compressors 24, 26. The latter comprise a plurality of rows of rotor blades associated with a plurality of rows of stator blades.

[0056] The rotor blades are driven integrally or kinematically by the shafts 30 , 34 about the axis of rotation 4 in order to generate an air flow and progressively compress the air flow to the inlet of the combustion chamber 22 .

[0057] The turbine includes a gear arrangement, such as a reduction gear 36, which drives the fan 18 at a lower rotational speed than the rotor blades of the low-pressure compressor. Thus, two turbines are sufficient to drive the compressor and the fan at three different respective rotational speeds.

[0058] In particular, the gearbox 36 allows the low pressure compressor 24 to be compact (fewer compression stages) because it can rotate at high speeds, particularly transonic speeds, while the blower 18 rotates at subsonic speeds.

[0059] exist Figure 1In the illustration of FIG, the low-pressure compressor 24 includes two annular rows of rotor blades 40, 42 that are integral with the shaft 34. Alternatively, the compressor may include three annular rows of rotor blades.

[0060] Compressor 24 includes an annular row of stator blades including an inlet row 39, an intermediate stator row 41 axially between two rotor rows, and a downstream stator row 45. Alternatively, when the compressor includes three rows of rotor blades, a second intermediate stator row is disposed between the second and third rotor rows.

[0061] The blades 39 may be variable stator vanes (VSV).In this application, the blades 39 will be considered to be in their angular position during flight ("cruise").

[0062] The compressor 24 is located directly downstream of a casing support 47 and directly upstream of another casing support 49 .

[0063] These housings 47, 49 may comprise an annular duct forming part of the primary flow duct 8 and may have struts 46, 48 extending radially through the flow duct 8. The annular duct of the housing may have a gooseneck profile. They may significantly reduce the diameter of the primary flow duct 8.

[0064] The upstream housing 47 and its arms 46 may support the gearbox 36 .

[0065] Upstream of the arm 46 is a row of stator blades 37 close to the nozzle 12. Downstream of the arm 48 is the high-pressure compressor 26 equipped with stator blades 50 and rotor blades 52. The latter are driven in rotation by the shaft 30.

[0066] Figure 2 Some parameters are shown that help define the geometry of the compression stage formed by the rotor R1 and the stator S1, viewed in a plane parallel to the axis 4 and perpendicular to one of these radii, at a distance r from the axis 4. More precisely, Figure 2 are some adjacent leaves on the plane ( Figure 3 The rotor blades 40 accelerate the fluid flow using the energy transmitted via the drive shaft 34. Due to the shape of the stator blades, the stator stages of the blades 41 "convert" the kinetic energy into pressure.

[0067] The blade rotates about the axis 4 at a rotational speed ω, which means that a point on the blade with radial coordinate r will have a speed U=rω. In the example shown, all rotor blades rotate at the same speed ω, so the speed in the cross section P is U for all rotor blades.

[0068] The camber lines of the blades 40 , 41 are labeled C40 and C41 and are shown in dashed lines.

[0069] The vector diagram to the left of the rotor blades 40 corresponds to the flow conditions at the compressor stage inlet. The fluid enters with an absolute velocity V1, depending on the direction of the turbine element preceding the compressor. The relative velocity vector W1 of the fluid entering the rotor is derived by subtracting the vector U from V1.

[0070] For comparison, another absolute inlet velocity, V1, is shown. V1' is more inclined relative to axis 4 than V1, but both velocities, V1 and V1', have the same axial component, Vx1. For the same rotational speed, U, it follows a relative velocity, W1', which is lower in absolute value than W1.

[0071] Therefore, for the same value of the axial component of the inlet velocity, the relative speed of the air with respect to the rotor is different if it is more or less inclined.

[0072] At the outlet of rotor R1, the fluid has a relative velocity W2, whose direction depends on the geometry of the blades. By adding vector U to vector W2, an absolute output velocity V2 is obtained, which is greater than V1.

[0073] The stator blades 41 then deflect and decelerate the fluid, this deceleration causing a pressure increase. The direction of the velocity V3 depends inter alia on the geometry of the stator blades.

[0074] The geometry of the blades can be described in part by their inlet and outlet angles.

[0075] The "inlet angle" of the blade is the angle formed at the leading edge in a plane parallel to the compressor axis and perpendicular to the radius between the camber tangent and the compressor axis.

[0076] The "exit angle" of a blade is the angle formed at the trailing edge of the blade in a plane parallel to the compressor axis and perpendicular to the radius between the tangent line of the camber and the compressor axis.

[0077] Thus, relative velocity W2 is generally oriented along the outlet angle of rotor blade 40, while velocity V3 is generally oriented along the outlet angle of stator blade 41. Under nominal conditions (blade tip speed between 360 m / s and 420 m / s), the difference between the outlet angle of blade 41 and flow direction V3 can be on the order of 5 or 6 degrees. The present invention is defined as a flow angle greater than 15 degrees on average and greater than 15 degrees between 60% and 80% of the blade height. It can also be defined by a stator blade outlet angle that is greater than 9 degrees on average and greater than 9 degrees in the range of 60-80%, preferably greater than 14 degrees.

[0078] According to Euler's equation, it can be noted that the enthalpy increase ΔH in the rotor R1 is U x (Vt2-Vt1), where U is the rotational speed of the rotor (in m / s), Vt1 and Vt2 are the tangential velocities of the air flow at the rotor inlet and outlet, respectively, in a reference frame fixed to the motor.

[0079] In the relative reference frame of the rotor, the tangential velocities of the air flow (at the inlet and outlet, respectively) are Wt1 = Vt1 - U and Wt2 = Vt2 - U.

[0080] If the axial velocities of the air flow at the rotor inlet and outlet are Vx1 and Vx2 respectively, the angles of the air flow relative to the rotor are α1 = arctan(Wt1 / Vx1) and α2 = arctan(Wt2 / Vx2), that is, α1 = arctan((Vt1-U) / Vx1) and α2 = arctan((Vt2-U) / Vx2).

[0081] Therefore, it is found that the enthalpy increase proportional to the tangential velocity difference between the rotor inlet and outlet (Vt2-Vt1) is not proportional but related to the rotor deflection α2-α1, which is actually arctan((Vt2-U) / Vx2)-arctan((Vt1-U) / Vx1).

[0082] The inlet air flow angle of an individual blade is recorded as 4x.1 and the outlet air flow angle is recorded as 4x.2 (where x = 1, 2, 3, 4 or 5).

[0083] It is worth noting that the angles 4x.1, 4x.2 vary over the height of the blade.

[0084] The difference between the inlet and outlet flow angles can be denoted as Δ4x = |4x.1–4x.2|. This difference also characterizes the camber of the blade. For example, Δ40 = |40.1–40.2|. If the angles differ significantly, the blade becomes highly cambered. For rotor blades, this results in significant aerodynamic loads. For stator blades, this can lead to significant flow deflection (deflection = the difference between the blade's inlet and outlet flow angles). Since the angles 4x.1 and 4x.2 vary along the height of the blade / vane, the arithmetic mean value over the entire blade / vane height is used and is denoted as <Δ4x>.

[0085] Load is the change in enthalpy divided by the square of the rotational speed. Therefore, a highly cambered blade will allow for greater compression. The camber limit is usually determined by mechanical strength, or in our case by the requirement to not choke at supersonic speeds.

[0086] According to the same principle as described above for V1′, the geometry of the blades influences the exit angle of the flow 41.2, which defines the direction of the velocity V3 and therefore the relative velocity value of the flow W3 relative to the rotor (not shown) downstream of S1. Thus, a larger value of the angle 41.2 means a smaller value of W3.

[0087] Figure 2 Also shown are the leading edges 40.3, 41.3, the trailing edges 40.4, 41.4, the pressure sides 40.5, 41.5 and the suction sides 40.6, 41.6 of the blades 40, 41.

[0088] Figure 3 An isometric view of a rotor blade 40 is shown, in particular showing a possible position of a plane P in which the Figure 2 The figure also shows the height H of the blade and the different parts of this height: H1 extends over 70% of the blade's height, on the inside; H2 extends over 30% of the blade, on the outside; H3 corresponds to the inner 40%; H4 to the inner 80% and H5 to the outer 20%.

[0089] Figure 3 The change in airflow angle 40.2 between the radial position of the blade root and the radial position of the blade head is also highlighted. The same is true for any angle 4x.1 or 4x.2 from the blade root (or "base") to the blade tip (or "head").

[0090] Figure 4 An isometric view of a stator blade 41 is shown. The figure shows the height H of the stator blade (substantially equal to the height of the preceding rotor blade) and the different parts of this height: H6 extends over 40% of the blade height, on the inside; H7 extends between 60% and 80% of the blade, on the outside; H8 corresponds to the inner 10%; H9 extends between the inner 10% and the outer 10%; and H10 corresponds to the outer 10%.

[0091] Figure 5 An example is shown in which the compressor comprises exactly two rows of rotor blades 40 , 44 and two rows of stator blades 41 , 45 .

[0092] According to the invention, the flow angle 41.2 is greater than 15° on average over the height H of the blade 41 and greater than 20° on average over the portion H7 forming the radial portion between 60% and 80% of the radial height of the blade 41. It will be understood that the geometry of the blade is smooth, without corners or angled points, providing, if necessary, a connecting radius between the portion greater than 15° and the portion greater than 20°.

[0093] Figure 5 The direction of the velocity V2 at the outlet of the blade 40 is also highlighted, which is achieved by the angle 40 . 2 being greater than 40°.

[0094] A difference in the angles Δ40 and Δ44 is also observed, the ratio Δ40 / Δ44 being greater than or equal to 1.4. As for the flow deflection Δ45 on the stator blade 45 , it is less than 50°.

[0095] Figure 6 An embodiment with three compression stages is shown. Figure 5 In comparison, the compressor comprises an additional row of rotor blades 42 and an additional row of intermediate stator blades 43 .

[0096] The ratio Δ40 / Δ44 is greater than or equal to 1.4, and the ratios Δ40 / Δ42 and Δ42 / Δ44 are greater than or equal to 1.2.

[0097] The angle 43.2 is greater than 15° on average over the height H of the blade 43 and extends radially outwardly over 60% to 80% of the radial height of the blade 43. Figure 4 H7) is greater than 15°, preferably 20°.

[0098] It should be noted that in this application, the rotor and stator blades have the same height H, but in practice there may be variations in the radial height and position of the blades.

[0099] Finally, all rotor blades here rotate at the same speed ω. A person skilled in the art will know how to apply the teaching of the present application to situations where the rotor blade rows have different speeds.

Claims

1. A transonic compressor for an aircraft turbine, comprising exactly two annular rows of rotor blades and exactly one annular row of stator blades between said rows of rotor blades, in, The rotor blades in the row of rotor blades each have a radial height and a radially outer end having a velocity, wherein the stator blades in the exactly one annular row of stator blades each have a radial height, wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves the row of stator blades in a direction forming a stator outlet air flow angle with the axial direction, Wherein, when the row of rotor blades is driven at a certain rotational speed so that the speed of the radial outer end of the rotor blade is between 360m / s and 420m / s, the average value of the stator outlet air flow angle along the stator blade height is greater than 15°, and the stator outlet air flow angle is greater than 15° at all distances to the rotation axis in the radial portion of the stator blade included between 60% and 80% of the radial height of the stator blade.

2. The transonic compressor according to claim 1, further comprising a downstream row of stator blades arranged downstream of exactly two rows of rotor blades, each stator blade in the downstream row having a radial height, in, In a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow enters the downstream row of stator blades in a direction forming a downstream stator inlet air flow angle with the axial direction, wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves the downstream row of stator blades in a direction forming a downstream stator outlet air flow angle with the axial direction, wherein, in a radial portion of the stator blades of the downstream row extending from 10% to 90% of their radial height, the difference between the downstream stator inlet angle and the downstream stator outlet angle is less than 40° for each distance from the axis of rotation, wherein, in a radial portion of the stator blade forming a radial inner portion of the stator blade less than 10% of its radial height, the difference between the downstream stator inlet angle and the downstream stator outlet angle is less than 50° for each distance to the axis of rotation, and Therein, in a radial portion of the stator blade forming a radial outer portion of the stator blade more than 90% of its radial height, the difference between the downstream stator inlet angle and the downstream stator outlet angle is less than 50° for each distance to the axis of rotation.

3. The transonic compressor according to claim 1, wherein: There are exactly two rows of rotor blades, including an upstream row and a downstream row. wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow enters each row of rotor blades along a direction forming a rotor inlet angle with the axial direction, wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves each row of rotor blades in a direction forming a rotor outlet angle with the axial direction, wherein the upstream row of rotor blades comprises an upstream average value along the height of the rotor blades of the difference between the rotor inlet angle and the rotor outlet angle for each given distance from the axis of rotation, wherein the downstream row of rotor blades comprises a downstream average along the height of the rotor blades of the difference between the rotor inlet angle and the rotor outlet angle for each given distance from the axis of rotation, and Among them, the upstream average value is greater than or equal to 1.4 times the downstream average value.

4. The transonic compressor according to claim 1, wherein: The rotor blade has a first inner portion extending from the inner end of the blade over 40% of the height of the rotor blade, a second inner portion extending from the inner end of the blade over 80% of the height of the rotor blade, and an outer portion extending from the outer end of the rotor blade over 20% of the height of the rotor blade, and wherein the rotor blade is designed so that the speed of the air flow relative to the rotor blade is less than or equal to Mach number 0.9 in the first inner portion, less than or equal to Mach number 1 in the second inner portion, and less than or equal to Mach number 1.05 in the outer portion.

5. The transonic compressor according to claim 1, in, In a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves each row of rotor blades in a direction forming a rotor outlet angle with the axial direction. The average value of the rotor outlet angle over the height of the rotor blade is greater than 40°.

6. The transonic compressor according to claim 1, wherein: Exactly two rows of rotor blades comprise an upstream row and a downstream row, wherein the speed at the exit of the rotor blades of the upstream row is below Mach number 0.

8.

7. The transonic compressor according to claim 1, wherein: Exactly two rows of rotor blades include an upstream row and a downstream row, and an enthalpy change of air flow between the outlet and inlet of the rotor blades in the upstream row is at least 1.4 greater than an enthalpy change of air flow between the outlet and inlet of the rotor blades in the downstream row.

8. A transonic compressor for an aircraft turbine engine, comprising exactly three annular rows of rotor blades and exactly two annular rows of intermediate stator blades between two consecutive rows of rotor blades, wherein the two rows of intermediate stator blades comprise an upstream row and a downstream row, in, The rotor blades in the row of rotor blades each have a radial height and a radially outer end having a velocity, wherein the stator blades in the row of middle stator blades each have a radial height, wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves the upstream row of stator blades in a direction forming a stator outlet air flow angle with the axial direction, Wherein, when the row of rotor blades is driven at a certain rotational speed so that the speed of the radial outer end of the rotor blade is between 360m / s and 420m / s, the average value of the stator outlet air flow angle along the stator blade height is greater than 15°, and the stator outlet air flow angle is greater than 15° at all distances to the rotation axis in the radial portion of the stator blade included between 60% and 80% of the radial height of the stator blade.

9. The transonic compressor according to claim 8, wherein: There are exactly three rows of rotor blades including upstream, middle and downstream rows. wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow enters each row of rotor blades along a direction forming a rotor inlet angle with the axial direction, wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves each row of rotor blades in a direction forming a rotor outlet angle with the axial direction, wherein the upstream row of rotor blades comprises an upstream average value along the height of the rotor blades of the difference between the rotor inlet angle and the rotor outlet angle for each given distance from the axis of rotation, wherein the middle row of rotor blades comprises a middle average along the height of the rotor blade of the difference between the rotor inlet angle and the rotor outlet angle for each given distance from the axis of rotation, and Among them, the upstream average is greater than or equal to 1.2 times the middle average.

10. The transonic compressor according to claim 9, wherein: The downstream row of rotor blades comprises a downstream average along the height of the rotor blade of the difference between the rotor inlet angle and the rotor outlet angle for each given distance from the axis of rotation, and Among them, the upstream average value is greater than or equal to 1.4 times the downstream average value.

11. The transonic compressor according to claim 9, wherein: The exactly three rows of rotor blades include an upstream row, a middle row, and a downstream row, the enthalpy change of the air flow between the outlet and the inlet of the rotor blades in the upstream row being at least 1.4 greater than the enthalpy change of the air flow between the outlet and the inlet of the rotor blades in the downstream row.

12. A turbomachine comprising a transonic low-pressure compressor, a high-pressure compressor, a structural casing upstream of the transonic low-pressure compressor, and an intermediate structural casing between the transonic low-pressure compressor and the high-pressure compressor, the structural casing defining a gooseneck-shaped air duct, wherein the transonic low-pressure compressor is arranged directly downstream of the upstream casing and directly upstream of the intermediate casing, and wherein the transonic low-pressure compressor comprises two annular rows of rotor blades and one annular row of stator blades between the rows of rotor blades, in, The rotor blades in the row of rotor blades each have a radial height and a radially outer end having a velocity, Wherein, the stator blades of the annular row of stator blades each have a radial height, wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves the row of stator blades in a direction forming a stator outlet air flow angle with the axial direction, Wherein, when the row of rotor blades is driven at a certain rotational speed so that the speed of the radial outer end of the rotor blade is between 360m / s and 420m / s, the average value of the stator outlet air flow angle along the stator blade height is greater than 15°, and the stator outlet air flow angle is greater than 15° at all distances to the rotation axis in the radial portion of the stator blade included between 60% and 80% of the radial height of the stator blade. 13 . The turbine according to claim 12 , further comprising a fan and a gearbox arranged between the transonic low-pressure compressor and the fan.

14. A method of operating a transonic compressor, wherein the transonic compressor comprises two annular rows of rotor blades and exactly one annular row of stator blades between the rows of rotor blades, in, The rotor blades in the row of rotor blades each have a radial height and a radially outer end having a velocity, wherein the stator blades in the exactly one annular row of stator blades each have a radial height, wherein, in a plane perpendicular to the radial direction and at a distance from the axis of rotation of the rotor blades, the air flow leaves the row of stator blades in a direction forming a stator outlet air flow angle with the axial direction, Wherein, when the row of rotor blades is driven at a certain rotational speed so that the speed of the radial outer end of the rotor blade is between 360m / s and 420m / s, the average value of the stator outlet air flow angle along the stator blade height is greater than 15°, and the stator outlet air flow angle is greater than 15° at all distances to the rotation axis in the radial portion of the stator blade included between 60% and 80% of the radial height of the stator blade.

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