A high pressure ratio transonic axial compressor

Abstract

The application discloses a high-pressure-ratio transonic axial flow compressor, which comprises a hub, front-row rotor blades, rear-row rotor blades and stator blades; the rear-row rotor blades are arranged between the front-row rotor blades and the stator blades; a plurality of front-row rotor blades and a plurality of rear-row rotor blades are uniformly arranged on the hub along a circumferential direction; and the front-row rotor blades and the rear-row rotor blades one-to-one correspond to form tandem rotors. The application realizes shock wave supercharging through the front-row rotor blades, and the rear-row rotor blades complete large airflow deflection; meanwhile, the application adopts an adsorption type stator to reduce or prevent airflow separation in a blade channel, thereby significantly improving a compressor stage pressure ratio and improving aerodynamic performance of the compressor. The application can realize a higher stage pressure ratio, and in the case of high inlet Mach number, high load and large airflow deflection angle, the application overcomes the problem that a high stage pressure ratio and efficiency are difficult to realize due to the fact that flow separation in the compressor cannot be controlled in the prior art.

Description

Technical Field

[0001] This invention belongs to the field of compressor technology, specifically a high-pressure ratio transonic axial flow compressor. Background Technology

[0002] Modern aircraft place higher demands on engine thrust-to-weight ratio. Currently, fourth-generation engines have achieved a thrust-to-weight ratio of 10. As one of the three core components of an aero-engine, reducing the number of stages and increasing the stage pressure ratio can effectively reduce engine size and weight, thereby improving the thrust-to-weight ratio. However, with the increase in compressor stage pressure ratio, flow separation and intensified secondary flow inevitably occur within the compressor. This leads to deterioration of internal flow, increased losses, and ultimately limits the improvement of stage pressure ratio. Therefore, controlling flow separation and further improving the compressor stage pressure ratio has become an important research topic in the aerodynamic design of aero-engine compressors.

[0003] Tandem blade technology can effectively control the velocity distribution and boundary layer development on the blade surface, offering advantages such as a large airflow turning angle and low total pressure loss. The principles behind tandem blade technology in controlling internal compressor flow and reducing flow losses are mainly: 1) The independent rows of blades allow the boundary layer to redevelop on the rear blades; 2) The gaps between the front and rear blade rows effectively blow away the boundary layer from the rear blades, thus inhibiting its development. Therefore, it can withstand higher blade loads without a significant increase in flow losses.

[0004] Boundary layer extraction technology, as a type of active flow control technology, has gained increasing attention from turbomachinery researchers since Professor Kerrebrock first proposed the concept of adsorption compressors in 1997, due to its ability to effectively control flow separation and increase blade load. The basic idea is to create slots (including holes or grooves) on the blade surface to remove low-energy fluid from the blade surface area, suppressing boundary layer separation and increasing the airflow turning angle; and to create holes or grooves along the flow direction or circumferentially on the casing and hub to remove low-velocity fluid accumulating in the endwall region, weakening the endwall effect and reducing secondary flow losses. Merchant designed a high-pressure-ratio adsorption compressor, achieving a design pressure ratio of 3.4 while simultaneously removing 2% of the inlet flow from the rotor and stator blade surfaces and a total of 3% from the casing and hub surfaces. However, while boundary layer extraction technology can effectively control flow separation, extraction within the rotor introduces structural challenges such as blade strength issues and extraction path arrangement. Summary of the Invention

[0005] To address the aforementioned problems, this invention discloses a high-pressure ratio transonic axial compressor, comprising: a hub, a front row of rotor blades, a rear row of rotor blades, and stator blades;

[0006] The rear rotor blades are positioned between the front rotor blades and the stator blades.

[0007] Multiple front-row rotor blades and multiple rear-row rotor blades are evenly arranged on the hub along the circumferential direction;

[0008] The front and rear rotor blades are arranged in a one-to-one correspondence to form a tandem rotor.

[0009] Furthermore, it also includes: the casing;

[0010] The front row of rotor blades, the rear row of rotor blades, and the stator blades are surrounded by an organic casing.

[0011] Furthermore, a suction slit is provided on the suction surface of the stator blade, and a hollow suction cavity is provided inside.

[0012] The suction slits are distributed along the spanwise direction of the stator blades;

[0013] The suction slit is connected to the hollow suction cavity.

[0014] Furthermore, it also includes: rectifier casing;

[0015] The stator blades are connected to the rectifier casing;

[0016] The rectifier casing has an annular hollow structure with an annular suction chamber inside;

[0017] The hollow suction cavities and the annular suction cavities are interconnected, and the hollow suction cavities and the annular suction cavities constitute a suction cavity;

[0018] The rectifier casing is connected to the casing.

[0019] Furthermore, it also includes: pipelines;

[0020] Multiple of the aforementioned pipelines are evenly arranged on the outside of the rectifier casing.

[0021] Furthermore, the suction slit is located at 50% to 60% of the chord length of the stator blade;

[0022] The suction rate of the multiple stator blades is 0.9–1.1 kg / s.

[0023] Furthermore, the chord length of the front rotor blades ranges from 91.9 to 112 mm, the inlet geometric angle ranges from 51 to 61°, and the outlet geometric angle ranges from 28 to 65°.

[0024] Furthermore, the chord length of the rear rotor blades ranges from 63 to 76 mm, the inlet geometric angle ranges from 35 to 61°, and the outlet geometric angle ranges from -24 to 42°.

[0025] Furthermore, the chord length of the stator blade ranges from 62.9 to 67 mm, the inlet geometric angle ranges from 44.1 to 45°, and the outlet geometric angle ranges from -1.5 to 0°.

[0026] Furthermore, the number of front and rear rotor blades is 30 to 45.

[0027] The number of the static blades is 70 to 94.

[0028] Compared with the prior art, the embodiments of the present invention have at least the following advantages:

[0029] 1) The front row of rotor blades in the tandem rotor has a high inlet Mach number, which enables shock wave pressurization; the channel normal shock wave is located at the trailing edge of the front row of rotor blades, which reduces the loss of mutual interference between the shock wave and the boundary layer.

[0030] 2) The Mach number decreases after passing through the front row of rotor blades, and pressure can be increased through a large bend angle in the rear row of rotor blades.

[0031] 3) By using independent blade rows, the boundary layer can redevelop on the rear rotor blades. At the same time, the gap between the front and rear blade rows can blow away the boundary layer on the rear rotor blades, thereby inhibiting the development of the boundary layer on the rear rotor blades. Therefore, it can withstand higher blade loads without significantly increasing flow losses.

[0032] 4) Since the rotor pressure ratio exceeds the conventional design, it will inevitably lead to a high stator inlet Mach number and a large airflow deflection angle, which greatly increases the stator load. By using adsorption stator blades, the internal flow separation of the stator can be controlled to achieve a low-loss design.

[0033] 5) Without the rotor blades drawing air, the compressor's total pressure ratio reaches 3.35, far exceeding the conventional load level, and the adiabatic efficiency reaches 87.1%.

[0034] Other features and advantages of the invention will be set forth in the following description, and will be apparent in part from the description, or may be learned by practicing the invention. The objects and other advantages of the invention can be realized and obtained by means of the structures pointed out in the description and the drawings. Attached Figure Description

[0035] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0036] Figure 1 A three-dimensional view of the high-pressure ratio transonic axial compressor of the present invention is shown;

[0037] Figure 2 An internal view of the high-pressure ratio transonic axial compressor of the present invention is shown;

[0038] Figure 3 A three-dimensional view of the high-pressure ratio transonic axial compressor of the present invention, viewed from the intake direction, is shown;

[0039] Figure 4 A three-dimensional view of the high-pressure ratio transonic axial compressor of the present invention, viewed from the outlet direction, is shown;

[0040] Figure 5 An axial two-dimensional cross-sectional view of the high-pressure ratio transonic axial compressor of the present invention is shown;

[0041] Figure 6 A schematic diagram of the adsorption-type stator blade structure is shown;

[0042] Figure 7 The blade profiles of the front row of rotor blades, the rear row of rotor blades, and the stator blades at the blade tip are shown.

[0043] Figure 8 The blade profile diagrams of the front row rotor blades, rear row rotor blades, and stator blades in the middle section of the blade are shown.

[0044] Figure 9 The blade profiles of the front row rotor blades, rear row rotor blades, and stator blades at the blade root are shown.

[0045] Figure 10 A high-load compressor meridional flow diagram according to an embodiment of the present invention is shown;

[0046] Figure 11 A schematic diagram of the velocity triangles of the front and rear rows of tandem rotors according to an embodiment of the present invention is shown;

[0047] Figure 12 Mach number cloud diagrams of the front and rear rotor blades at 90% blade height according to an embodiment of the present invention are shown.

[0048] Figure 13 Mach number cloud diagrams of the front and rear rotor blades at 50% blade height according to an embodiment of the present invention are shown.

[0049] Figure 14 Mach number cloud diagrams of the front and rear rotor blades at 10% blade height according to an embodiment of the present invention are shown.

[0050] Figure 15 A comparison diagram of the flow field before and after suction at the root of the stator blade according to an embodiment of the present invention is shown;

[0051] Figure 16 A total pressure ratio contour plot of the compressor outlet section according to an embodiment of the present invention is shown;

[0052] In the diagram: 1. Hub; 2. Casing; 3. Front rotor blades; 4. Rear rotor blades; 5. Stator blades; 6. Rectifier casing; 7. Suction slot; 8. Suction chamber; 9. Piping. Detailed Implementation

[0053] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0054] Under the premise of high pressure ratio design, this invention adopts a tandem blade design for the rotor to avoid problems such as structural strength caused by boundary layer suction in rotating components. The front rotor blades 3 perform shock wave pressurization, and the rear rotor blades 4 complete a large airflow deflection, which greatly improves the rotor's working capacity. At the same time, the adsorption-type stator blades 5 reduce or prevent airflow separation in the blade channel, which significantly improves the compressor stage pressure ratio and improves the compressor's aerodynamic performance. The compressor speed is 10431~11530 r / min, the flow rate is 65.8~72.8 kg / s, the pressure ratio reaches 3.35~3.5, and the adiabatic efficiency is not less than 87.1%.

[0055] like Figure 1 As shown, the present invention proposes a high-pressure ratio transonic axial compressor, comprising: a hub 1, a casing 2, front rotor blades 3, rear rotor blades 4, and stator blades 5;

[0056] The rear rotor blades 4 are disposed between the front rotor blades 3 and the stator blades 5.

[0057] Multiple front-row rotor blades 3 (e.g.) Figure 3 As shown in the figure, multiple rear rotor blades 4 are evenly arranged on the hub 1 along the circumferential direction;

[0058] Hub 1 is used to drive the front rotor blades 3 and the rear rotor blades 4 to rotate at the same speed.

[0059] The front rotor blades 3, the rear rotor blades 4 and the stator blades 5 are surrounded by an organic box 2;

[0060] The casing 2 is used to house the front row of rotor blades 3, the rear row of rotor blades 4, and the stator blades 5.

[0061] The front rotor blades 3 and the rear rotor blades 4 are arranged in a tandem rotor, which is used to significantly increase the blade load without significantly increasing the flow loss.

[0062] The front row of rotor blades 3 of the tandem rotor has a high inlet Mach number, which is used to achieve shock wave pressurization. After passing through the front row of rotor blades 3, the Mach number decreases, and the rear row of rotor blades 4 is used for large-angle pressurization. Through the independent front row of rotor blades 3 and rear row of rotor blades 4, the boundary layer redevelops on the rear row of rotor blades 4. At the same time, the gap between the front and rear rows of blades has the function of blowing away the boundary layer of the rear row of rotor blades 4, thereby inhibiting the development of the boundary layer of the rear row of rotor blades 4.

[0063] The stator blade 5 is used for airflow rectification and deceleration / boosting.

[0064] like Figure 2 As shown, a suction slit 7 is provided on the suction surface of the stator blade 5, and a hollow suction cavity is provided inside (a hollow suction cavity is opened between the suction surface and the pressure surface). A rectifier casing 6 and a pipeline 9 are provided outside the stator blade 5.

[0065] The suction slit 7 is distributed along the spanwise direction of the stator blade 5;

[0066] The suction slit 7 is connected to the hollow suction cavity.

[0067] Suction slit 7 is used to remove low-energy fluid from the suction surface, thereby weakening or eliminating flow separation;

[0068] like Figure 5 As shown, the stator blade 5 is not connected to the hub 1, but is connected to the rectifier casing 6;

[0069] Figure 6 (a) is a schematic diagram of the stator blade structure. For example... Figure 6 As shown in (a), the rectifier casing 6 is an annular hollow structure with an annular suction cavity inside; for example, the cross-section of the annular suction cavity is rectangular, rounded rectangle, polygonal, elliptical, etc.

[0070] Figure 6 (b) is a cross-sectional view of the stator blade. For example... Figure 6 As shown in (b), the multiple hollow suction cavities and the annular suction cavity are connected to each other, and the multiple hollow suction cavities and the annular suction cavity form a suction cavity 8;

[0071] The rectifier casing 6 is welded or bolted to the casing 2.

[0072] The suction chamber 8 is used to contain low-energy fluid drawn in through the suction slit 7.

[0073] The rectifier casing 6 is used to enclose the suction chamber 8 and rectify the low-energy fluid.

[0074] Multiple of the aforementioned pipes 9 are evenly arranged on the outside of the rectifier casing 6.

[0075] Pipeline 9 is used to discharge low-energy fluids out of the system through suction slot 7 and suction chamber 8.

[0076] In some embodiments, the inlet radius of the hub 1 ranges from 228 to 252.5 mm. Preferably, the inlet radius of the hub 1 is 240.4 mm.

[0077] The inlet radius of the casing 2 is in the range of 415–459 mm. Preferably, the inlet radius of the casing 2 is 434.17 mm.

[0078] In some embodiments, the suction slit 7 is located at 50% to 60% of the chord length of the stator blade 5; preferably, the suction slit 7 is located at 55% of the chord length of the stator blade 5, which has the most significant effect on controlling the flow separation of the stator.

[0079] The suction rate of the multiple stator blades 5 is 0.9–1.1 kg / s. Preferably, the suction rate of the multiple stator blades 5 is 1.0 kg / s, which has the most significant effect on controlling the flow separation of stators.

[0080] In some embodiments, the chord length of the front rotor blades 3 ranges from 91.9 to 112 mm, the inlet geometric angle ranges from 51 to 61°, and the outlet geometric angle ranges from 28 to 65°.

[0081] In some embodiments, the chord length of the rear rotor blades 4 ranges from 63 to 76 mm, the inlet geometric angle ranges from 35 to 61°, and the outlet geometric angle ranges from -24 to 42°.

[0082] In some embodiments, the chord length of the stator blade 5 ranges from 62.9 to 67 mm, the inlet geometry angle ranges from 44.1 to 45°, and the outlet geometry angle ranges from -1.5 to 0°.

[0083] In some embodiments, the number of front rotor blades 3 and rear rotor blades 4 is 30 to 45; preferably, the number of front rotor blades 3 and rear rotor blades 4 is 37.

[0084] The number of stator blades 5 is 70 to 94. Preferably, the number of stator blades 5 is 82.

[0085] Within the range of chord length, inlet geometry angle, outlet geometry angle, and number of blades mentioned above, the compressor can achieve optimal aerodynamic performance.

[0086] The high-pressure transonic axial compressor rotor blades have a higher inlet Mach number, achieving shock wave pressurization in the front rotor blades 3, with the blade tip shock wave located at the trailing edge of the front rotor blades 3. After passing through the front rotor blades 3, the Mach number decreases, allowing for pressurization through a large bend angle in the rear rotor blades 4 channel. The independent blade rows allow the boundary layer to regrow on the rear rotor blades 4. The gaps between the front and rear blade rows have the function of blowing away the boundary layer on the rear rotor blades 4, thereby inhibiting the development of the boundary layer on the rear rotor blades 4. Therefore, it can withstand higher blade loads without a significant increase in flow losses.

[0087] Since the loads on the front rotor blades 3 and the rear rotor blades 4 exceed those of conventional designs, it is inevitable that the stator blades 5 will have a high inlet Mach number and a large airflow deflection angle, which greatly increases the stator load. By using adsorption-type stator blades 5, the internal flow separation of the stator can be controlled, and a low-loss design can be achieved.

[0088] Example 1

[0089] like Figure 3 As shown, there are 37 rotor blades in the front row 3 and 37 rotor blades in the rear row 4, which are evenly distributed on the hub 1 along the circumferential direction.

[0090] like Figure 4 As shown, there are a total of 82 stator blades 5; the low-energy fluid is discharged from the external pipeline 9 through the suction slit 7 and the suction chamber 8.

[0091] The high-pressure transonic compressor of this invention has a flow rate of 69.3 kg / s, a pressure ratio of 3.35, and an adiabatic efficiency of 87.1%. The specific parameters are shown in Table 1. The flow separation phenomenon inside the compressor is well controlled. Figure 10 A meridional flow diagram of a high-load compressor according to an embodiment of the present invention is shown. Figure 10 As shown, the front rotor blades 3 and the rear rotor blades 4 are two rows of tandem rotors, and the rear stator blades 5 perform boundary layer suction at the 55% chord length position.

[0092] Table 1 Parameters of High Pressure Ratio Transonic Compressors

[0093]

[0094]

[0095] The chord lengths, inlet geometric angles, outlet geometric angles, and relative positions of the blade roots, mid-sections, and tips of the front and rear rotor blades are shown in Table 2. Figure 7 , Figure 8 , Figure 9 The leaf shape diagrams are shown at three sections: leaf tip, leaf middle, and leaf root.

[0096] Table 2 Geometric parameters of transonic compressor rotor and stator for high pressure ratio.

[0097] parameter leaf tips In the leaves Leaf roots Front rotor chord length / mm 112 91.9 92 Front rotor inlet geometry angle / ° 61 55 51 Front rotor outlet geometry angle / ° 65 50 28 Rear rotor chord length / mm 76 63 63 Rear rotor inlet geometry angle / ° 61 45 35 Rear rotor outlet geometry angle / ° 42 6 -24 Rotor axial overlap / mm 15.2 15.2 15.2 Rotor circumferential position / mm 13.5 10.5 7.1 Stator chord length / mm 67 62.9 63.5 Stator inlet geometry angle / ° 45 44.1 45 Stator exit geometry angle / ° -0.5 -1.5 0

[0098] Figure 11 The velocity triangles of the front and rear tandem rotors according to an embodiment of the present invention are shown. Figure 11 (a) is the velocity triangle of the front rotor. Figure 11 (b) is the velocity triangle of the rear rotor. For example... Figure 11 As shown in (a), the front rotor undergoes shock wave pressurization, with the relative velocity direction remaining essentially unchanged, while the axial velocity decreases significantly. Here, C1 is the inlet absolute velocity of the front rotor blade 3; C2 is the outlet absolute velocity of the front rotor blade 3; C 1a C represents the inlet axial velocity of the front rotor blade 3. 2a U1 is the exit axial velocity of the front rotor blade 3; U2 is the inlet circumferential velocity of the front rotor blade 3; W1 is the inlet relative velocity of the front rotor blade 3; W2 is the outlet relative velocity of the front rotor blade 3. For example... Figure 11 As shown in (b), the Mach number at the inlet of the rear rotor is lower, allowing for pressurization through a larger rotation angle. Here, C2 represents the absolute inlet velocity of the rear rotor blade 4, and its value is equal to... Figure 11 In (a), C2 and C3 represent the absolute exit velocity of the rear rotor blades 4; U2 represents the inlet circumferential velocity of the rear rotor blades 4, which is equal to... Figure 11 In (a), U2 and U3 are the exit circumferential velocities of the rear rotor blades 4; W2 is the inlet relative velocity of the rear rotor blades 4, and its value is equal to... Figure 11 In (a), W2 and W3 are the relative exit velocities of the rear rotor blades 4.

[0099] Because the absolute airflow angle at the outlet of the rear rotor blade 4 is too large, a large airflow deflection needs to be completed in the stator. Therefore, boundary layer suction is required in the stator. A suction slit 7 runs through the blade height at 55% chord length of the stator. The blade interior is a hollow suction cavity. Low-energy fluid enters the blade cavity through the suction slit 7 and is discharged through the suction slit 7 and suction cavity 8 into the external pipe 9. Figure 6 As shown.

[0100] Depend on Figure 12 , 13As shown in section 14, the front rotor blades 3 form shock wave structures of varying intensities from root to tip. The tip section of the front rotor blades 3 exhibits a typical double shock wave structure: an oblique shock wave and a channel normal shock wave. The normal shock wave is located at the trailing edge exit of the front rotor blades 3, preventing flow separation induced by interference between the shock wave and the boundary layer. Due to the high inlet Mach number of the front rotor blades 3 and their small blade bend angle, the flow relies mainly on shock wave pressurization. After passing through the front rotor blades 3, the Mach number decreases, allowing the bend angle of the rear rotor blades 4 to increase, primarily relying on the large bend angle for pressurization. Furthermore, the independent blade rows allow the boundary layer to redevelop on the rear rotor blades 4. Simultaneously, the gap between the front and rear blade rows helps to remove the boundary layer from the rear rotor blades 4, thus inhibiting its development. Therefore, the overall flow inside the rotor is relatively smooth.

[0101] Figure 15 A comparison diagram of the flow field before and after suction at the root of the stator blades according to an embodiment of the present invention is shown. Figure 15 As shown in (a), the flow field before suction at 5-10% of the stator blade height is observed. Due to the high Mach number of 1.26 at the inlet of stator blade 5 and the blade curvature reaching 45 degrees, a large-scale flow separation occurs on the back of the stator blade at 10% of the blade height before suction, extending from approximately 45% of the chord length to the channel outlet. The circumferential separation range accounts for one-quarter of the channel. Figure 15 (b) shows the flow field after the stator blade was pumped at 510% of its height. The flow was significantly improved after pumping, and the separation scale in both the chord and circumferential directions was significantly reduced.

[0102] Figure 16 A total pressure ratio contour plot of the compressor outlet section according to an embodiment of the present invention is shown. Figure 16 It can be seen that the total pressure ratio of the compressor with low blade span is very high, reaching about 3.8, and the total pressure ratio of the entire compressor is also as high as 3.35.

[0103] The high-pressure ratio transonic axial compressor proposed in this invention has at least the following advantages:

[0104] 1) The front row of rotor blades in the tandem rotor has a high inlet Mach number, which enables shock wave pressurization; the channel normal shock wave is located at the trailing edge of the front row of rotor blades, which reduces the loss of mutual interference between the shock wave and the boundary layer.

[0105] 2) The Mach number decreases after passing through the front row of rotor blades, and pressure can be increased through a large bend angle in the rear row of rotor blades.

[0106] 3) By using independent blade rows, the boundary layer can redevelop on the rear rotor blades. At the same time, the gap between the front and rear blade rows can blow away the boundary layer on the rear rotor blades, thereby inhibiting the development of the boundary layer on the rear rotor blades. Therefore, it can withstand higher blade loads without significantly increasing flow losses.

[0107] 4) Since the rotor pressure ratio exceeds the conventional design, it will inevitably lead to a high stator inlet Mach number and a large airflow deflection angle, which greatly increases the stator load. By using adsorption stator blades, the internal flow separation of the stator can be controlled to achieve a low-loss design.

[0108] 5) Without the rotor blades drawing air, the compressor's total pressure ratio is 3.35, far exceeding the conventional load level, and the adiabatic efficiency reaches 87.1%.

[0109] Although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A high-pressure ratio transonic axial-flow compressor, characterized in that, include: Hub (1), front rotor blades (3), rear rotor blades (4) and stator blades (5); The rear rotor blades (4) are disposed between the front rotor blades (3) and the stator blades (5); Multiple front rotor blades (3) and multiple rear rotor blades (4) are evenly arranged on the hub (1) along the circumferential direction; The front row rotor blades (3) and the rear row rotor blades (4) are arranged in a one-to-one correspondence to form a tandem rotor; The static blade (5) has a suction slit (7) on its suction surface and a hollow suction cavity inside. The suction slit (7) is distributed along the spanwise direction of the stator blade (5); The suction slit (7) is connected to the hollow suction cavity; Also includes: casing (2); The front rotor blades (3), the rear rotor blades (4) and the stator blades (5) are surrounded by a casing (2); Also includes: rectifier casing (6); The stator blade (5) is connected to the rectifier casing (6); The rectifier casing (6) has an annular hollow structure and an annular suction chamber inside; The hollow suction chambers and the annular suction chamber are connected to each other, and the hollow suction chambers and the annular suction chambers form a suction chamber (8). The rectifier casing (6) is connected to the casing (2); The suction slit (7) is located at 50%~60% of the chord length of the stator blade (5); The suction capacity of the multiple stator blades (5) is 0.9~1.1 kg / s.

2. The high-pressure ratio transonic axial compressor according to claim 1, characterized in that, Also includes: Pipeline (9); Multiple of the aforementioned pipes (9) are evenly arranged on the outside of the rectifier casing (6).

3. The high-pressure ratio transonic axial flow compressor according to claim 1, characterized in that, The chord length of the front rotor blades (3) ranges from 91.9 to 112 mm, the inlet geometric angle ranges from 51 to 61°, and the outlet geometric angle ranges from 28 to 65°.

4. The high-pressure ratio transonic axial compressor according to claim 1, characterized in that, The chord length of the rear rotor blades (4) ranges from 63 to 76 mm, the inlet geometric angle ranges from 35 to 61°, and the outlet geometric angle ranges from -24 to 42°.

5. The high-pressure ratio transonic axial flow compressor according to claim 1, characterized in that, The chord length of the stator blade (5) ranges from 62.9 to 67 mm, the inlet geometric angle ranges from 44.1 to 45°, and the outlet geometric angle ranges from -1.5 to 0°.

6. The high-pressure ratio transonic axial compressor according to claim 1, characterized in that, The number of the front rotor blades (3) and the rear rotor blades (4) is 30 to 45. The number of the static blades (5) is 70 to 94.

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