Gas turbine compressor and control method

Abstract

The application discloses a gas turbine compressor, comprising: a barrel-shaped casing; a hub installed in the casing; a rotor rotatably installed on the hub, the rotor comprising a plurality of rotor blades distributed in a circumferential direction outside the hub, an outer wall of the casing being provided with a plurality of air injection portions in the circumferential direction, each air injection portion being in communication with an inside of the casing and being arranged close to a leading edge of a tip of the rotor blade and being configured to inject air towards the rotor blade to increase a flow speed of air flow along an axial direction of the tip of the rotor blade; and a plurality of stator blades arranged in the circumferential direction on the hub, each stator blade being configured to inject air along a wall of a blade body of the stator blade to weaken flow separation of air flow passing through a suction surface of the stator blade.

Description

Technical Field

[0001] At least one embodiment of the present invention relates to the field of axial flow compressor technology, specifically to a gas turbine compressor and a control method thereof. Background Technology

[0002] Stable and efficient operation under wide operating conditions is the development requirement and inevitable trend of future aero-engine and gas turbine compressor power equipment. As its core unit, the compression system has a significant impact on the performance of aero-engine and gas turbine compressors in terms of pressure ratio, efficiency, and stable operating margin. Seeking a compression system that can balance efficiency and stability is the core technical problem in the field of aero-engine and gas turbine compressor aerodynamic stability.

[0003] Existing gas turbine compressors typically have only a single structure for jetting air towards the top of the rotor blades or only a single stator blade for jetting air along its own blade body. Therefore, existing gas turbine compressors can only either increase the stability margin of the gas turbine compressor or only improve its efficiency. Consequently, existing gas turbine compressors cannot simultaneously increase both the stability margin and efficiency, failing to meet the operational requirements of gas turbines under conditions such as start-up, low-speed cruise, and variable operating conditions. Summary of the Invention

[0004] In view of this, the present invention provides a gas turbine compressor, which, by providing a plurality of jet outlets circumferentially towards the rotor blades on the outer wall of the casing and by configuring the stator blades to jet along their own blade walls, can simultaneously increase the flow velocity of the airflow along the axial direction of the rotor blade tip and reduce the flow separation of the airflow flowing through the suction surface of the stator blades.

[0005] According to an embodiment of the present invention, a gas turbine compressor is provided, comprising: a cylindrical casing; a hub installed within the casing; a rotor rotatably mounted on the hub, the rotor including a plurality of rotor blades distributed circumferentially on the outer side of the hub, a plurality of jet nozzles disposed circumferentially on the outer wall of the casing, each jet nozzle communicating with the interior of the casing and disposed near the leading edge of the rotor blade tip, and configured to jet towards the rotor blade to increase the flow velocity of the airflow flowing axially along the blade tip of the rotor blade; and a plurality of stator blades disposed circumferentially on the hub, each stator blade being configured to jet along its own blade wall to reduce flow separation of the airflow flowing through the suction surface of the stator blade.

[0006] According to an embodiment of the present invention, the jet section is configured as an arc-shaped jet channel, which extends radially along the housing and is inclined circumferentially along the housing. An inlet and an outlet are respectively provided at both ends of the arc-shaped jet channel along the radial direction of the housing, so that external airflow flows into the interior of the arc-shaped jet channel through the inlet and is sprayed onto the tip of the rotor blade through the outlet. The side of the arc-shaped jet channel that is inclined circumferentially along the housing and extends longitudinally is an arc surface.

[0007] According to an embodiment of the present invention, the cross section of the jet section is formed with a first side and a second side facing each other. The first side includes a first straight segment extending in the radial direction of the housing and a first arc segment with its center located inside the cross section, which are connected in sequence. The second side includes a second straight segment extending in the radial direction of the housing, a second arc segment with its center located inside the cross section and a third arc segment with its center located outside the cross section, which are connected in sequence.

[0008] According to an embodiment of the present invention, the first arc segment is tangent to the first straight segment and the housing, the second arc segment is tangent to the second straight segment, and the two ends of the third arc segment are tangent to the second arc segment and the housing, respectively.

[0009] According to an embodiment of the present invention, the distance L between the end of the first side of the jet section and the leading edge of the rotor blade tip is 5 times the distance between the rotor blade tip and the inner wall of the housing; the distance W between the end of the first side and the end of the second side of the jet section is 8 times the distance between the rotor blade tip and the inner wall of the housing; and / or, the number of jet sections is half the number of rotor blades.

[0010] According to an embodiment of the present invention, the stator blade includes: a stator blade body, wherein the suction surface of the stator blade body sequentially includes a first suction surface and a second suction surface along the direction from the leading edge of the stator blade body to the trailing edge of the stator blade body; an air intake channel formed along the radial direction of the housing on the inside of the stator blade body near the leading edge of the stator blade body; and an air jet slit extending along the radial direction of the housing and formed on the inside of the stator blade body near the trailing edge of the stator blade body, wherein the air jet slit communicates with the air intake channel, so that external airflow flows into the interior of the stator blade body through the air intake channel, and flows sequentially through the air jet slit and the air jet nozzle to the second suction surface, and flows to the trailing edge of the stator blade body.

[0011] According to an embodiment of the present invention, the length of the jet slit along the circumferential direction of the housing is 0.2-0.5 mm.

[0012] According to an embodiment of the present invention, the second suction surface is tangent to the first suction surface, and the inner wall of the jet slit on the side of the suction surface near the stator body is connected to the portion of the second suction surface of the stator body near the leading edge of the stator body by an arc; wherein, the second suction surface is approximately formed as an arc surface with a radius of 31.08 mm.

[0013] According to an embodiment of the present invention, a control method for a gas turbine compressor is also provided, applicable to the gas turbine compressor described in the above embodiments. The control method includes: obtaining an operating condition corresponding to the incoming flow speed and the inlet flow rate of the gas turbine compressor inlet using a stability enhancement model; when the operating condition is a stable operating condition, obtaining a first mass flow rate of the stator blade jet corresponding to the incoming flow speed and the inlet flow rate using the stability enhancement model, and jetting along its own blade body at the first mass flow rate to reduce flow separation of the airflow on the suction surface of the stator blade; and when the operating condition is a near-stall point operating condition, obtaining a second mass flow rate of the jet jet and a third mass flow rate of the jet jet of the stator blade using the stability enhancement model, such that the jet jet jets towards the blade tip of the rotor blade at the second mass flow rate, and the stator blade jets along its own blade body at the third mass flow rate, to increase the flow velocity of the airflow flowing in the axial direction along the blade tip of the rotor blade, while reducing flow separation of the airflow on the suction surface of the stator blade.

[0014] According to an embodiment of the present invention, the enhanced stability and efficiency model is obtained through the following steps: Based on the incoming airflow speed at the gas turbine compressor inlet, the inlet flow rate of the gas turbine compressor inlet, the mass flow rate of the stator blade jet, and the mass flow rate of the jet from the jet section, the corresponding compressor efficiency and stable operating margin of the gas turbine compressor are calculated to obtain multiple sets of values ​​for compressor efficiency and stable operating margin corresponding to different incoming airflow speeds, inlet flow rates, stator blade jet mass flow rates, and jet section mass flow rates; based on the multiple sets of values, a enhanced stability and efficiency model is trained using a neural network algorithm to achieve the acquisition of operating conditions corresponding to the incoming airflow speed and inlet flow rate, as well as the mass flow rates of the stator blade jet and the jet section jet corresponding to the incoming airflow speed and inlet flow rate.

[0015] According to the above embodiments of the present invention, the gas turbine compressor, by providing multiple jet outlets that jet towards the rotor blades along the circumferential direction on the outer wall of the casing, and by configuring the stator blades to jet along their own blade walls, simultaneously increases the flow velocity of the airflow flowing along the axial direction of the rotor blade tips and reduces the flow separation of the airflow flowing through the suction surface of the stator blades. This allows for both widening the stable operating margin of the gas turbine compressor and improving its operating efficiency, thereby meeting the operational needs of the gas turbine under conditions such as start-up, low-speed cruise, and variable operating conditions. Attached Figure Description

[0016] Figure 1 This is a schematic diagram of the assembly between the rotor blades, jet section and stator blades of a gas turbine compressor according to an embodiment of the present invention.

[0017] Figure 2 This is an assembly diagram of the casing, hub, rotor, and jet section of a gas turbine compressor according to an embodiment of the present invention.

[0018] Figure 3 This is a schematic diagram of the assembly between the jet section and rotor blades of a gas turbine compressor according to an embodiment of the present invention;

[0019] Figure 4 This is a schematic diagram of the jet injection of the stator blades of the gas turbine compressor according to an embodiment of the present invention;

[0020] Figure 5 This is a schematic diagram of the assembly of the jet section and the casing of a gas turbine compressor according to an embodiment of the present invention;

[0021] Figure 6 This is a top view of the assembly of the jet section and the casing of a gas turbine compressor according to an embodiment of the present invention;

[0022] Figure 7 This is a plan view of the assembly of the jet section, casing, and rotor blades of a gas turbine compressor according to an embodiment of the present invention;

[0023] Figure 8 This is a plan view of the assembly dimensions between the jet section, casing, and rotor blades of a gas turbine compressor according to an embodiment of the present invention.

[0024] Figure 9 This is a three-dimensional schematic diagram of the stator blades of a gas turbine compressor according to an embodiment of the present invention;

[0025] Figure 10 yes Figure 9 A magnified view of part A in the middle;

[0026] Figure 11 This is a system diagram of a control method for a gas turbine compressor according to an embodiment of the present invention;

[0027] Figure 12 This is a flowchart of a control method for a gas turbine compressor according to an embodiment of the present invention; and

[0028] Figure 13 This is a schematic diagram illustrating the effect of stabilization and efficiency enhancement of the gas turbine compressor in an embodiment of the present invention.

[0029] In the picture:

[0030] 1-Casing;

[0031] 2-Wheel hub;

[0032] 3-Rotor;

[0033] 31-Rotor blade; 311-Leading edge of blade tip; 312-Blade tip;

[0034] 4-Jet section;

[0035] 41-Arc-shaped jet passage;

[0036] 411 - Input port;

[0037] 412 - Output port;

[0038] 413 - First side; 4131 - First straight line segment; 4132 - First circular arc segment;

[0039] 414 - Second side; 4141 - Second straight line segment; 4142 - Second circular arc segment; 4143 - Third circular arc segment;

[0040] 5-Stationary blades;

[0041] 51-Leaf body;

[0042] 52 - Suction surface; 521 - First suction surface; 522 - Second suction surface;

[0043] 53-Still leaf body; 531-Leading edge; 532-Tail edge;

[0044] 54 - Air intake channel;

[0045] 55-Air jet seam;

[0046] 56 - Jet nozzle;

[0047] 6-Air intake;

[0048] 7-Air outlet;

[0049] 8-Guide vane. Detailed Implementation

[0050] To make the objectives, technical solutions, and advantages of the present invention clearer, the present invention will be further described in detail below with reference to specific embodiments and accompanying drawings.

[0051] According to one aspect of the invention, a gas turbine compressor is provided, comprising: a cylindrical casing; a hub installed within the casing; a rotor rotatably mounted on the hub, the rotor including a plurality of rotor blades distributed circumferentially on the outer side of the hub, a plurality of jet nozzles disposed circumferentially on the outer wall of the casing, each jet nozzle communicating with the interior of the casing and disposed near the leading edge of the rotor blade tip, and configured to jet towards the rotor blade to increase the flow velocity of the airflow flowing axially along the blade tip of the rotor blade; and a plurality of stator blades disposed circumferentially on the hub, each stator blade being configured to jet along its own blade wall to reduce flow separation of the airflow flowing through the suction surface of the stator blade.

[0052] Figure 1 This is a schematic diagram of the assembly between the rotor blades, jet section and stator blades of a gas turbine compressor according to an embodiment of the present invention. Figure 2 This is an assembly diagram of the casing, hub, rotor, and jet section of a gas turbine compressor according to an embodiment of the present invention. Figure 3 This is a schematic diagram of the assembly between the jet section and rotor blades of a gas turbine compressor according to an embodiment of the present invention; Figure 4 This is a schematic diagram of the jet injection of the stator blades of the gas turbine compressor according to an embodiment of the present invention; Figure 5 This is a schematic diagram of the assembly of the jet section and the casing of a gas turbine compressor according to an embodiment of the present invention; Figure 6 This is a top view of the assembly of the jet section and the casing of a gas turbine compressor according to an embodiment of the present invention; Figure 7 This is a plan view of the assembly of the jet section, casing, and rotor blades of a gas turbine compressor according to an embodiment of the present invention.

[0053] According to an exemplary embodiment of the present invention, please refer to Figures 1-7 A gas turbine compressor is provided, comprising a cylindrical casing 1, a hub 2, a rotor 3, and a plurality of stator blades 5. The hub 2 is mounted inside the casing 1. The rotor 3 is rotatably mounted on the hub 2. The rotor 3 includes a plurality of rotor blades 31 distributed circumferentially on the outer side of the hub 2. A plurality of jet nozzles 4 are provided circumferentially on the outer wall of the casing 1, each jet nozzle 4 communicating with the interior of the casing 1 and located near the leading edge 311 of the rotor blade 31, and configured to jet towards the rotor blade 31 to increase the flow velocity of the airflow flowing along the axial direction of the blade tip 312 of the rotor blade 31. The plurality of stator blades 5 are arranged circumferentially on the hub 2, and each stator blade 5 is configured to jet along its own blade body 51 attached to the wall to reduce flow separation of the airflow flowing through the suction surface 52 of the stator blade 5.

[0054] In this embodiment, by providing multiple jet nozzles 4 along the circumferential direction on the outer wall of the casing 1 to jet towards the rotor blades 31, and by configuring the stator blades 5 to jet along their own blade body 51, the flow velocity of the airflow flowing along the axial direction of the blade tip 312 of the rotor blades 31 is increased, and the flow separation of the airflow flowing through the suction surface 52 of the stator blades 5 is reduced. This allows for both widening the stable operating margin of the gas turbine compressor and improving the working efficiency of the gas turbine compressor, so as to meet the operating needs of the gas turbine under conditions such as start-up, low-speed cruise, and variable operating conditions.

[0055] In this embodiment, the housing 1 is provided with multiple openings (not shown in the figure), and the output port 412 of each jet 4 is correspondingly provided with one opening, so that the multiple jet 4 communicate with the interior of the housing 1 through the multiple openings provided on the housing 1, so as to realize jetting to the tip 312 of the rotor blade 31. The jet 4 are evenly arranged along the circumferential direction of the housing 1.

[0056] The outlet 412 of the jet section 4 adopts a Coanda effect nozzle structure to form an airflow that flows along the blade tip 312 of the rotor blade 31.

[0057] In some exemplary embodiments, reference is made to Figure 1 , Figure 3 as well as Figures 5-7 The jet section 4 is constructed as an arc-shaped jet passage 41. The arc-shaped jet passage 41 extends radially along the housing 1 and is inclined circumferentially along the housing 1. An inlet 411 and an outlet 412 are respectively provided at both ends of the arc-shaped jet passage 41 along the radial direction of the housing 1, so that external airflow flows into the arc-shaped jet passage 41 through the inlet 411 and is sprayed onto the tip 312 of the rotor blade 31 through the outlet 412. The longitudinally extending side of the arc-shaped jet passage 41, which is inclined circumferentially along the housing 1, is an arc surface.

[0058] With the above configuration, external airflow flows into the arc-shaped jet channel 41 through the inlet 411 and is injected onto the wall of the blade tip 312 of the rotor blade 31 through the outlet 412. This increases the flow velocity of the airflow along the axial direction of the blade tip 312 of the rotor blade 31, enhances the flow momentum of the airflow from the rotor blade 31 to the stator blade 5 along the hub 2, delays the occurrence of gas turbine compressor stall, and widens the stable operating margin of the gas turbine compressor.

[0059] Figure 8 This is a plan view showing the assembly dimensions between the jet section, casing, and rotor blades of a gas turbine compressor according to an embodiment of the present invention.

[0060] In some exemplary embodiments, reference is made to Figures 5-8The jet section 4 has a cross-section with a first side 413 and a second side 414 facing each other. The first side 413 includes a first straight segment 4131 extending radially along the housing 1 and a first arc segment 4132 whose center is located inside the cross-section, which are connected in sequence. The second side 414 includes a second straight segment 4141 extending radially along the housing 1, a second arc segment 4142 whose center is located inside the cross-section, and a third arc segment 4143 whose center is located outside the cross-section, which are connected in sequence.

[0061] In some exemplary embodiments, reference is made to Figure 8 The first arc segment 4132 is tangent to the first straight segment 4131 and the housing 1 respectively, the second arc segment 4142 is tangent to the second straight segment 4141, and the two ends of the third arc segment 4143 are tangent to the second arc segment 4142 and the housing 1 respectively.

[0062] It should be noted that, in this embodiment, the radius of the first arc segment 4132 is R1, the radius of the second arc segment 4142 is R3, and the radius of the third arc segment 4143 is R2. The radius R3 of the second arc segment 4142 and the radius R2 of the third arc segment 4143 are approximately equal, while the radius R2 of the third arc segment 4143 is greater than the radius R1 of the first arc segment 4132.

[0063] In some exemplary embodiments, reference is made to Figures 1-3 as well as Figure 8 The distance L between the end of the first side 413 of the jet section 4 and the leading edge 311 of the rotor blade 31 is 5 times the distance between the tip 312 of the rotor blade 31 and the inner wall of the housing 1. The distance W between the end of the first side 413 and the end of the second side 414 of the jet section 4 is 8 times the distance between the tip 312 of the rotor blade 31 and the inner wall of the housing 1.

[0064] In this embodiment, by optimizing the shape of the first side 413 and the second side 414 of the cross section of the jet section 4, and optimizing the distance L between the end of the first side 413 of the jet section 4 and the leading edge 311 of the rotor blade 31 and the spacing W between the end of the first side 413 and the end of the second side 414 of the jet section 4, the output port 412 of the jet section 4 is made to have a Coanda effect nozzle structure.

[0065] In this embodiment, the Coanda Effect refers to the tendency of a fluid (water or air) to deviate from its original flow direction and instead flow along a convex surface. When there is surface friction between the fluid and the surface it flows over (also known as fluid viscosity), the fluid will flow along the surface of the object as long as the curvature is not too large. According to Newton's third law, if an object exerts a deflecting force on a fluid, the fluid will also exert a deflecting force on the object in the opposite direction.

[0066] Reference Figures 1-2 The number of jet nozzles 4 is half the number of rotor blades 31. Specifically, there are 8 jet nozzles 4 and 16 rotor blades 31, meaning that there is one jet nozzle 4 for every two rotor blades 31.

[0067] Figure 9 This is a three-dimensional schematic diagram of the stator blades of a gas turbine compressor according to an embodiment of the present invention; Figure 10 yes Figure 9 A magnified view of part A in the middle.

[0068] In some exemplary embodiments, reference is made to Figure 4 as well as Figures 9-10 The stator blade 5 includes a stator body 53, an air intake channel 54, and an air jet slit 55. The suction surface 52 of the stator body 53, along the direction from the leading edge 531 to the trailing edge 532 of the stator body 53, sequentially includes a first suction surface 521 and a second suction surface 522. The air intake channel 54 is formed radially inside the stator body 53 near the leading edge 531. The air jet slit 55 extends radially inside the stator body 53 near the trailing edge 532. The air jet slit 55 communicates with the air intake channel 54, allowing external airflow to flow into the stator body 53 through the air intake channel 54, and then sequentially through the air jet slit 55 and the air jet nozzle 56 to the second suction surface 522, and finally to the trailing edge 532 of the stator body 53.

[0069] With the above configuration, the external airflow flows into the interior of the stator blade body 53 through the air intake channel 54, and then flows sequentially through the jet slit 55 and the jet nozzle 56 to the second suction surface 522, and then to the trailing edge 532 of the stator blade body 53. This allows the stator blade 5 to spray air along its own blade body 51, thereby increasing the momentum of the low-energy airflow in the suction surface 52 of the stator blade 5, enhancing its ability to resist the reverse pressure gradient in the internal passage of the gas turbine compressor, and thus reducing the flow separation of the airflow flowing through the suction surface 52 of the stator blade 5, thereby improving the efficiency of the gas turbine compressor.

[0070] The jet slits 55 are arranged parallel to each other at different axial positions on the suction surface 52 of the stator blade 5. The jet slits 55 are usually opened before the position where flow separation occurs on the stator blade 5.

[0071] In detail, the jet slit 55 can be provided on the top of the stator blade 5 along the direction from the leading edge 531 of the stator blade 5 to the trailing edge 532 of the stator blade 5, wherein the end of the jet slit 55 near the trailing edge 532 of the stator blade 5 is usually opened before the position where flow separation occurs on the stator blade 5.

[0072] In some exemplary embodiments, reference is made to Figures 9-10 The length of the jet slit 55 along the circumferential direction of the casing 1 is 0.2-0.5 mm.

[0073] In this embodiment, the distance of the jet jet along the wall of the stator blade 5 remains consistent throughout the entire stator blade 5. Considering machining factors, the length of the jet jet slit 55 along the circumferential direction of the housing 1 is 0.2-0.5 mm.

[0074] In some exemplary embodiments, reference is made to Figures 9-10 The second suction surface 522 is tangent to the first suction surface 521. The inner wall of the jet slit 55 on the side of the suction surface 52 near the stator body 53 is connected to the part of the second suction surface 522 near the leading edge 531 of the stator body 53 by an arc. The second suction surface 522 is approximately formed as an arc surface with a radius of 31.08 mm.

[0075] In this embodiment, the second suction surface 522 is formed by connecting 20 arc surfaces in sequence, wherein the average radius of the 20 arc surfaces is 31.08 mm.

[0076] It should be noted that, in this embodiment, the jet section 4 and the stator blades 5 are respectively connected to an external air source or the downstream extraction pipeline of the gas turbine compressor. The casing 1 is respectively equipped with a first valve (not shown in the figure) for controlling the connection between the jet section 4 and the external air source, and a second valve (not shown in the figure) for controlling the connection between the interior of the stator blades 5 and the external air source. By controlling the opening and closing of the first and second valves respectively, the jet section 4 and the stator blades 5 are controlled to eject air outwards. The first and second valves are proportional solenoid valves.

[0077] Figure 11 This is a system diagram of a control method for a gas turbine compressor according to an embodiment of the present invention; Figure 12 This is a flowchart of a control method for a gas turbine compressor according to an embodiment of the present invention; Figure 13 This is a schematic diagram illustrating the effect of stabilization and efficiency enhancement of the gas turbine compressor in an embodiment of the present invention.

[0078] According to an exemplary embodiment of the present invention, please refer to Figures 11-13 A control method for a gas turbine compressor is provided, applicable to the gas turbine compressor described in the above embodiments. The control method includes: obtaining the operating condition corresponding to the incoming flow speed and inlet flow rate at the gas turbine compressor inlet 6 using a stability enhancement model, based on the incoming flow speed and inlet flow rate at the gas turbine compressor inlet 6. When the operating condition is stable, obtaining the first mass flow rate of the jet ejected from the stator blade 5 corresponding to the incoming flow speed and inlet flow rate using the stability enhancement model, and ejecting the jet along its own blade body 51 at the first mass flow rate to reduce flow separation of the airflow on the suction surface 52 of the stator blade 5. Under near-stall conditions, the second mass flow rate of the jet from the jet 4 and the third mass flow rate of the jet from the stator blade 5 are obtained through the stability enhancement model. This allows the jet 4 to jet towards the tip 312 of the rotor blade 31 with the second mass flow rate, and the stator blade 5 to jet along its own blade body 51 with the third mass flow rate. This increases the flow velocity of the airflow along the axial direction of the tip 312 of the rotor blade 31, while reducing the flow separation of the airflow on the suction surface 52 of the stator blade 5.

[0079] By using the above control method, the mass flow rate of the jet from the stator blade 5 and the jet from the jet section 4 can be controlled separately under different operating conditions. This avoids the mixing loss caused by the misalignment of the airflow from the jet section 4 and the stator blade 5 with the mainstream airflow velocity direction in the gas turbine compressor. It also avoids the weakening of the flow effect of the jet towards the blade tip 312 of the rotor blade 31 due to changes in the incoming speed and inlet flow rate of the airflow at the gas turbine compressor inlet 6. This can further reduce losses, improve the efficiency of the gas turbine compressor, and broaden the stable operating range.

[0080] Furthermore, when the gas turbine compressor is running, the efficiency of the gas turbine compressor can be improved by jetting air along the wall of the stator blade 5 via the stator blade 5. The stable operating margin of the gas turbine compressor is widened by jetting air onto the tip 312 of the rotor blade 31 via the jetting section 4. However, different incoming speeds and inlet flow rates at the gas turbine compressor inlet 6 typically correspond to different mass flow rates of the jetting air from the stator blade 5 and the jetting air from the jetting section 4. Changing the incoming speed and inlet flow rate requires real-time adjustment of the mass flow rates of the jetting air from the stator blade 5 and the jetting air from the jetting section 4; otherwise, the jetting speeds of the stator blade 5 and the jetting air from the jetting section 4 will differ too much from the mainstream airflow speed of the gas turbine compressor, increasing mixing losses and ultimately degrading the flow state of the gas turbine compressor. By adjusting the mass flow rate of the jet from the stator blade 5 and the jet from the jet section 4 in real time, it can be ensured that the jet volume of the stator blade 5 and the jet section 4 is the optimal jet volume under all operating conditions, thus ensuring that the gas turbine compressor achieves maximum efficiency and operates in a high-efficiency state at all times.

[0081] It should be noted that, in this embodiment, the gas turbine compressor is provided with an air inlet 6 and an air outlet 7 at both ends extending along the axial direction. The air inlet 6 is located on the side near the rotor 3, and the air outlet 7 is located on the side near the stator blade 5. The rotor 3 is also provided with a guide vane 8 for guiding the airflow on the side near the air inlet 6.

[0082] In some exemplary embodiments, reference is made to Figures 11-13 The expansion, stabilization, and efficiency enhancement model is obtained through the following steps:

[0083] Based on the incoming speed of the airflow at the inlet 6 of the gas turbine compressor, the inlet flow rate of the airflow at the inlet 6 of the gas turbine compressor, the mass flow rate of the jet from the stator blade 5, and the mass flow rate of the jet from the jet section 4, the compressor efficiency and stable operating margin of the gas turbine compressor are calculated to obtain multiple sets of values ​​for compressor efficiency and stable operating margin corresponding to different incoming speeds, inlet flow rates, mass flow rates of the jet from the stator blade 5, and mass flow rates of the jet from the jet section 4.

[0084] Based on these multiple sets of values, a stability-enhancing and efficiency-improving model is trained using a neural network algorithm. This model enables the acquisition of the operating conditions corresponding to the incoming flow speed and inlet flow rate, as well as the mass flow rate of the stator blade 5 and the mass flow rate of the jet from the jet section 4, which are also corresponding to the incoming flow speed and inlet flow rate.

[0085] To ensure the efficient and safe operation of gas turbine compressors, researchers have studied various flow control methods to improve compressor efficiency and extend their stable operating range. Tip-mounted jetting along the rotor blades offers the advantage of increasing compressor stall margin without reducing or increasing efficiency. This can be achieved by improving airflow at the blade tips through high-speed jetting, thus delaying rotor stall. Researchers such as Suder et al. conducted detailed numerical simulations and experimental verifications of steady-state jetting at the rotor blade tips, demonstrating that tip-mounted jetting effectively widens the stable operating range across the entire speed range of the gas turbine compressor. Around 2000, the team led by Nie Chaoqun at the Institute of Engineering Thermophysics, Chinese Academy of Sciences, conducted experimental research on tip-mounted micro-jet stabilization control methods for two-stage and three-stage low-speed axial compressors. With a jet flow rate of only 0.045%-0.056% of the gas turbine compressor's design flow rate, they were able to widen the compressor's stable operating boundary by 5.83%. While jet injection is being applied to the tips of rotor blades, many scholars both domestically and internationally have confirmed that wall-mounted jet flow separation control technology can enable fluid to flow along the suction surface of the stator blades, reducing flow separation and improving compressor efficiency. Researchers Hill et al. optimized the inlet guide vanes of a three-stage gas turbine compressor at Wright-Patterson Air Force Base, using the minimum jet flow rate and pressure loss as the objective function. Their research shows that inlet guide vanes utilizing the Coanda Effect have good performance. Fischer from the University of Hanover, Germany, applied wall-mounted jet injection to the stator blade design of a high-speed axial compressor, analyzing the impact of active flow control on the three-dimensional flow field. The results show that using a 1% jet flow rate can reduce the overall loss coefficient by approximately 21%. In recent years, the team led by Du Juan at the Institute of Engineering Thermophysics, Chinese Academy of Sciences, has verified the feasibility of wall-mounted jet injection in suppressing flow separation through experiments and simulations. The jetting of the rotor blade tip along the wall can improve the stable operating margin of the gas turbine compressor, while the jetting of the stator blade along its own blade wall can improve the compressor efficiency.

[0086] This embodiment of the gas turbine compressor, combining jet injection from the jet section 4 towards the tip 312 of the rotor blade 31 and jet injection from the stator blade 5 along its own blade body 51, provides a feasible technical approach for the gas turbine compressor to operate stably and efficiently under a wider range of operating conditions. Simultaneously, by recognizing the differences in the incoming flow speed and inlet flow rate at the gas turbine compressor inlet 6, the mass flow rate of the jet injection from the stator blade 5 and the jet injection from the jet section 4 are adjusted in real time to improve the overall performance of the gas turbine under conditions such as start-up, low-speed cruise, and variable operating conditions. This has significant practical value and scientific significance.

[0087] The specific embodiments described above further illustrate the purpose, technical solution, and beneficial effects of the present invention. It should be understood that the above descriptions are merely specific embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A gas turbine compressor, comprising: Cylindrical housing (1); The hub (2) is installed inside the housing (1); The rotor (3) is rotatably mounted on the hub (2). The rotor (3) includes a plurality of rotor blades (31) distributed in the circumferential direction on the outer side of the hub (2). The outer wall of the housing (1) is provided with a plurality of jet nozzles (4) in the circumferential direction. Each jet nozzle (4) communicates with the interior of the housing (1) and is located near the leading edge (311) of the blade tip of the rotor blade (31). It is configured to jet towards the rotor blade (31) to increase the flow velocity of the airflow along the axial direction of the blade tip (312) of the rotor blade (31). The cross section of the jet nozzle (4) includes a first arc segment (4132) with the center located inside the cross section, a second arc segment (4142) and a third arc segment (4143) with the center located outside the cross section. The jet nozzle (4) is configured as an output port (412) with a Coanda effect nozzle structure. as well as Multiple stator blades (5) are arranged on the hub (2) in the circumferential direction. Each stator blade (5) is configured to spray air along its own blade body (51) to reduce the flow separation of the airflow flowing through the suction surface (52) of the stator blade (5). Wherein, the distance L between the end of the first arc segment (4132) and the leading edge (311) of the rotor blade (31) is 5 times the distance between the tip (312) of the rotor blade (31) and the inner wall of the housing (1); the distance W between the end of the first arc segment (4132) and the end of the third arc segment (4143) is 8 times the distance between the tip (312) of the rotor blade (31) and the inner wall of the housing (1).

2. The gas turbine compressor according to claim 1, wherein, The jet section (4) is configured as an arc-shaped jet channel (41), which extends along the radial direction of the housing (1) and is inclined along the circumferential direction of the housing (1). The arc-shaped jet channel (41) is provided with an inlet (411) and an outlet (412) at both ends along the radial direction of the housing (1), so that the external airflow flows into the interior of the arc-shaped jet channel (41) through the inlet (411) and is sprayed onto the tip (312) of the rotor blade (31) through the outlet (412). The side of the arc-shaped jet channel (41) that is inclined in the circumferential direction along the housing (1) and extends in the longitudinal direction is an arc surface.

3. The gas turbine compressor according to claim 2, wherein, The cross-section of the jet section (4) has a first side (413) and a second side (414) facing each other. The first side (413) includes a first straight segment (4131) and a first arc segment (4132) extending in the radial direction of the housing (1) in sequence, which are smoothly connected in sequence. The second side (414) includes a second straight segment (4141), a second arc segment (4142) and a third arc segment (4143) extending in the radial direction of the housing (1) in sequence, which are smoothly connected in sequence.

4. The gas turbine compressor according to claim 3, wherein, The first arc segment (4132) is tangent to the first straight segment (4131) and the housing (1) respectively, the second arc segment (4142) is tangent to the second straight segment (4141), and the two ends of the third arc segment (4143) are tangent to the second arc segment (4142) and the housing (1) respectively.

5. The gas turbine compressor according to claim 3, and / or, the number of jet units (4) is half the number of rotor blades (31).

6. The gas turbine compressor according to claim 1, wherein, The stator blade (5) includes: The still blade body (53) has a suction surface (52) that includes a first suction surface (521) and a second suction surface (522) in sequence along the direction from the leading edge (531) of the still blade body (53) to the trailing edge (532) of the still blade body (53). An air intake channel (54) is formed along the radial direction of the housing (1) on one side of the stator vane body (53) near the leading edge (531) of the stator vane body (53); and The jet slit (55) extends radially along the housing (1) and is formed inside the stator body (53) on one side near the trailing edge (532) of the stator body (53). The jet slit (55) is connected to the air intake channel (54), so that the external airflow flows into the interior of the stator body (53) through the air intake channel (54), and flows sequentially through the jet slit (55) and the jet nozzle (56) to the second suction surface (522), and flows to the trailing edge (532) of the stator body (53).

7. The gas turbine compressor according to claim 6, wherein, The length of the jet slit (55) along the circumferential direction of the housing (1) is 0.2-0.5 mm.

8. The gas turbine compressor according to claim 6, wherein, The second suction surface (522) is tangent to the first suction surface (521), and the inner wall of the jet slit (55) on the side of the suction surface (52) near the stationary blade body (53) is connected to the part of the second suction surface (522) of the stationary blade body (53) near the leading edge (531) of the stationary blade body (53) by an arc. The second suction surface (522) is roughly formed as an arc surface with a radius of 31.08 mm.

9. A control method for a gas turbine compressor, applicable to the gas turbine compressor according to any one of claims 1-8, the control method comprising: Based on the incoming speed of the airflow at the gas turbine compressor inlet (6) and the inlet flow rate at the gas turbine compressor inlet (6), the operating conditions corresponding to the incoming speed and the inlet flow rate are obtained through the stability enhancement model. When the operating condition is stable, the first mass flow rate of the stator blade (5) corresponding to the incoming flow speed and the inlet flow rate is obtained by the stability enhancement model, and the first mass flow rate is used to spray air along its own blade body (51) to reduce the flow separation of the airflow on the suction surface (52) of the stator blade (5); and when the operating condition is near the stall point, the second mass flow rate of the jet (4) and the third mass flow rate of the jet (5) are obtained by the stability enhancement model, so that the jet (4) sprays air towards the tip (312) of the rotor blade (31) with the second mass flow rate, and the stator blade (5) sprays air along its own blade body (51) with the third mass flow rate, so as to increase the flow velocity of the airflow along the axial direction of the tip (312) of the rotor blade (31), and at the same time reduce the flow separation of the airflow on the suction surface (52) of the stator blade (5).

10. The control method for a gas turbine compressor according to claim 9, wherein, The stability enhancement and efficiency improvement model is obtained through the following steps: Based on the incoming speed of the airflow at the gas turbine compressor inlet (6), the inlet flow rate of the airflow at the gas turbine compressor inlet (6), the mass flow rate of the jet from the stator blades (5), and the mass flow rate of the jet from the jet section (4), the compressor efficiency and stable operating margin of the gas turbine compressor are calculated to obtain multiple sets of values ​​of compressor efficiency and stable operating margin corresponding to different incoming speeds, inlet flow rates, mass flow rates of the jet from the stator blades (5), and mass flow rates of the jet from the jet section (4). Based on the multiple sets of values, a stability enhancement model is trained using a neural network algorithm to achieve the following: by inputting the incoming flow speed and inlet flow rate, the working conditions corresponding to the incoming flow speed and inlet flow rate are obtained, as well as the mass flow rate of the stator blade (5) and the mass flow rate of the jet from the jet section (4) corresponding to the incoming flow speed and inlet flow rate.

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