Turbine guider

Abstract

The invention belongs to the technical field of aero-engine technologies, and provides a turbine guider which comprises a plurality of connected blade assemblies, each blade assembly comprises a plurality of blades, a first margin plate and a second margin plate, one end of each blade is connected with the corresponding first margin plate, the other end of each blade is connected with the corresponding second margin plate, and the first margin plates and the second margin plates are concentrically arranged; the axis on the surface formed by connecting each blade and the second edge plate is parallel to the axial line of the engine, and the angle formed between the tangent line on the first edge plate and the radial height of the blade is gradually reduced until the angle is gradually increased at the preset position of the blade; and the area of a channel formed between the adjacent blades is gradually reduced until the area of the channel is gradually increased after the blades are located at the preset positions of the blades. The channel area is controlled to change according to the setting, the transverse pressure gradient of the front portion of the channel between the blades is reduced, the radial pressure gradient of the blades is reduced, the pressure distribution of the end wall area of the blades is improved, the end wall surface load of the blades is changed, and the turbine efficiency is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of aeroengines, and in particular relates to a turbine guide vane. Background Art

[0002] At present, the development trend of aerodynamic design of turbine components of aircraft engines mainly has two aspects. One of them is the continuous improvement of the overall cycle parameters of turbine components, such as the continuous increase of turbine inlet temperature. This way of continuously increasing the turbine inlet temperature is strongly dependent on advanced aerodynamic and cooling design, as well as materials and manufacturing processes. With the increase of turbine inlet temperature, the coupling matching design of high-pressure turbine and combustion chamber has attracted more and more attention. The distribution characteristics of aero-thermal parameters such as hot spots, strong swirl and high turbulence at the combustion chamber outlet directly affect the stagnation distinction offline deviation and the vortex structure of secondary flow and channel vortex of the downstream high-pressure turbine stage. At the same time, the migration characteristics of hot spots and swirls in the turbine blade channel affect its aerodynamic loss, heat transfer performance and cooling characteristics. Therefore, it is necessary to study the aero-thermal performance of the high-pressure turbine stage under the inlet conditions of non-uniform aero-thermal parameters under the interaction between the combustion chamber and the turbine. However, the existing high-pressure turbine guide vane does not consider the influence of non-uniform aero-thermal parameter distribution such as hot spots, strong swirl and high turbulence at the combustion chamber outlet on the turbine aero-thermal performance when designing. When the flow field and temperature field at the combustion chamber outlet are uneven, it is difficult for the blades to adapt to the strong pre-swirl flow, and the blade profile loss is large, resulting in large cold air mixing losses, low turbine efficiency, poor cooling effect and easy ablation. Summary of the invention

[0003] In view of the above problems, the present invention proposes a turbine guide vane, comprising: a plurality of connected blade assemblies, each of the blade assemblies comprising a plurality of blades, a first edge plate and a second edge plate, one end of the plurality of blades being connected to the first edge plate, and the other end being connected to the second edge plate, the first edge plate and the second edge plate being arranged concentrically, the axis on the surface formed by the connection between each blade and the second edge plate being parallel to the axial line of the engine, the angle formed by the tangent on the first edge plate and the radial height of the blade gradually decreases until the angle gradually increases after the preset position of the blade, so that the channel area formed between adjacent blades decreases from large to small until the channel area increases from small to large after the preset position of the blade.

[0004] Optionally, the inlet configuration angle of the plurality of blades is 110°-130°.

[0005] Optionally, the outlet angles of the plurality of blades gradually decrease.

[0006] Optionally, the outlet angle of the plurality of blades is 20°-30°.

[0007] Optionally, the installation angle of the blade is 40°-50°.

[0008] Optionally, the thickness of the front half of the cross section of the blade remains constant along the chord length direction or varies within a preset thickness range along the chord length direction.

[0009] Optionally, each of the blades is installed at a different angle between the first edge plate and the second edge plate, so that a plurality of groups of flow channels with different areas are formed between the blades.

[0010] Optionally, a trailing edge slit is provided at the exhaust end of the blade, a pressure surface trailing edge point is provided at one end of the trailing edge slit, and a suction surface trailing edge point is provided at the other end, and the pressure surface trailing edge point and the suction surface trailing edge point are connected by ribs forming the trailing edge slit.

[0011] Optionally, air film holes are provided on the blade, the first edge plate and the second edge plate.

[0012] Optionally, the number of blades in each blade assembly is 4 to 8.

[0013] The turbine guide of the present invention has an axis on the surface formed by the connection between each blade and the second edge plate that is parallel to the axial line of the engine, and an angle formed by the tangent on the first edge plate and the radial height of the blade gradually decreases until the angle gradually increases after the preset position of the blade, so that the channel area formed between adjacent blades decreases from large to small until the channel area increases from small to large after the preset position of the blade, thereby controlling the channel area to change according to the setting, thereby reducing the lateral pressure gradient in the front part of the channel between the blades and reducing the radial pressure gradient of the blades, thereby improving the pressure distribution in the blade end wall area, changing the end wall surface load, and improving the turbine efficiency.

[0014] Other features and advantages of the present invention will be described in the following description, and partly become apparent from the description, or understood by practicing the present invention. The purpose and other advantages of the present invention can be realized and obtained by the structures pointed out in the description, claims and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0016] Figure 1 A schematic longitudinal cross-sectional view of one of the blades of the turbine guide device in an embodiment of the present invention is shown;

[0017] Figure 2 A three-dimensional schematic diagram of a blade of a turbine guide device in an embodiment of the present invention is shown;

[0018] Figure 3 A schematic diagram of a blade inlet passage of a turbine guide device in an embodiment of the present invention is shown;

[0019] Figure 4 A schematic diagram of a blade passage outlet of a turbine guide device in an embodiment of the present invention is shown;

[0020] Figure 5 A schematic cross-sectional view of a turbine guide device in an embodiment of the present invention is shown;

[0021] Figure 6 A schematic diagram of cross-sectional parameters of a turbine guide vane in an embodiment of the present invention is shown.

[0022] In the figure, 11, blade; 12, first edge plate; 13, second edge plate; 16, boss; 2, casing plate; 5, cross section; 50, trailing edge split; 51, trailing edge point of pressure surface; 52, trailing edge point of suction surface; 110, sealing groove. DETAILED DESCRIPTION

[0023] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0024] like Figure 1As shown, the present invention provides a turbine guide vane, comprising: a plurality of connected blade assemblies, each blade assembly comprising a plurality of blades 11, a first edge plate 12 and a second edge plate 13, a cascade being arranged at the top position of the blade 11, one end of the plurality of blades 11 being connected to the first edge plate 12, and the other end being connected to the second edge plate 13, the first edge plate 12 and the second edge plate 13 being arranged coaxially, the axis on the surface formed by the connection between each of the blades 11 and the second edge plate 13 being parallel to the axial line of the engine, the angle formed between the tangent on the first edge plate 12 and the radial height of the blade 11 gradually decreases until the angle gradually increases after the preset position of the blade 11, the preset position being at the blade end position, the channel structure between the blades being basically of a tapered type, the tapering angle being from 30° to 25°, so that the channel area formed between adjacent blades 11 is from large to small until the channel area is from small to large after the preset position of the blade 11. Thus, by controlling the channel area to change according to the setting, the lateral pressure gradient in the front of the channel between the blades can be reduced, that is, the lateral pressure gradient in the front of the cascade channel on the blade can be reduced, and the radial pressure gradient of the suction surface on the blade 11 is reduced, thereby improving the pressure distribution in the end wall area of ​​the blade 11, changing the surface load of the cascade at its end wall, reducing the secondary flow loss of the end wall, and improving the turbine efficiency. It should be noted that the two flow channel designs can determine the height and axial size of the engine steering shaft. Optionally, the inner side surface of the first edge plate 12 is a revolving surface, thereby realizing the surface matching with the blade, and the second edge plate 13 is an arc surface, a revolving surface or a circular surface.

[0025] like Figure 6 As shown, in one embodiment, the inlet structural angle of the plurality of blades 11 is 110°-130°. The inlet structural angle of 110°-130° can adapt to the swirl at the combustion chamber outlet, wherein the swirl angle is 30°-40°. The leading edge pressure surface and the suction surface of the inlet structural angle are basically symmetrically designed, and the attack angle has good adaptability, which is convenient for adapting to the swirl at the combustion chamber outlet, so that the swirl at the combustion chamber outlet can pass vertically through the channel of the blade 11.

[0026] like Figure 4 As shown, in one embodiment, the outlet angles of the plurality of blades 11 gradually decrease. It should be noted that the outlet angle of the blade 11 refers to the angle between the tangent of the blade outlet edge and the tangent of the intersection point of the outer circle of the impeller (first edge plate 12). The outlet angle of the blade 11 is designed with a reverse linear distribution law that is large at the root and small at the top. This distribution form is conducive to increasing the reaction force at the root of the blade 11, making the blade cascade groove on the blade 11 have good convergence, and also making the root of the blade 11 where the rotor is located have good acceleration.

[0027] like Figure 6As shown, in one embodiment, the outlet angle of the plurality of blades 11 is 20°-30°. By setting a suitable outlet angle range, the aerodynamic load at the tip of a high reaction force turbine can be reduced, and the tip leakage loss can be reduced.

[0028] In one embodiment, the installation angle of the blade 11 is 40°-50°. It should be noted that installing the blade 11 within the range of 40°-50° is beneficial to improving the cooling effect of the structure.

[0029] like Figure 4 As shown, in one embodiment, the thickness of the front half of the cross section 5 of the blade 11 remains unchanged along the chord length direction or changes within a preset thickness range along the chord length direction. The cross section 5 can also be called a blade profile, and the relative maximum thickness of the blade profile remains unchanged along the radial direction. It should be noted that the chord length is formed by the line connecting the bottom of the blade profile and the top tangent of the blade profile. The axial chord length of the blade 11 remains basically unchanged along the radial direction, and the installation angle of the blade 11 changes slightly along the radial direction, which is close to the design of a straight blade 11. Moreover, the blade profile is short and thick as a whole (short in the horizontal direction). Specifically, the maximum relative thickness of the blade profile remains unchanged along the radial direction, and the thickness of the blade profile in the front half of the blade profile does not change much along the chord direction. This design is conducive to the design of complex cooling structures and improves the cooling effect to adapt to high-temperature incoming flow conditions.

[0030] like Figure 6 As shown, in one embodiment, the blades 11 are installed at different angles between the first edge plate 12 and the second edge plate 13, so that several groups of flow channels with different areas are formed between the blades 11. It should be noted that this installation angle is equivalent to the installation angle. Specifically, the blades 11 are divided into several groups according to the different blade installation angles, and guides with different flow areas are matched by changing the number of combinations of different groups to meet the performance matching needs of the whole machine. The area difference of different groups is between ±3%. The present invention can adapt well to the design of the combustion chamber, and is convenient for matching and adaptive adjustment of the performance of the whole machine.

[0031] like Figure 2 As shown, in one embodiment, a trailing edge split 50 is provided at the exhaust end of the blade 11, and a pressure surface trailing edge point 51 is provided at one end of the trailing edge split 50, and a suction surface trailing edge point 52 is provided at the other end, and the pressure surface trailing edge point 51 and the suction surface trailing edge point 52 are connected by ribs forming the trailing edge split 50. It should be noted that the trailing edge split 50 is composed of a plurality of split ribs arranged at intervals and in parallel, such as Figure 5As shown, the design of the split rib strips of the trailing edge (exhaust end) of the blade 11 is different from the conventional design. The present invention directly connects the pressure surface trailing edge point 51 of the blade with the suction surface trailing edge point 52 of the blade. This design further reduces the thickness of the trailing edge, reduces the blockage and loss of the trailing edge, improves the cooling effect, and improves the flow capacity and performance. In addition, the trailing edge of the present invention is set at the minimum channel area, while the conventional trailing edge may be set at a larger channel area and a more forward position, and the pressure surface trailing edge point 51 is not connected to the suction surface trailing edge point 52, so the cooling effect is not as good as that of the present invention, and the flow capacity is not as strong. In addition, the pressure surface trailing edge point 51 and the suction surface trailing edge point 52 in the present invention are located at the outlet or exhaust edge of the blade 11, wherein the trailing edge split is also at the outlet or exhaust edge of the blade 11, the pressure surface trailing edge point 51 is on the inner side of the trailing edge split, and the suction surface trailing edge point 52 is on the outer side of the trailing edge split, and the pressure surface trailing edge point 51 is connected to the suction surface trailing edge point 52 through the trailing edge split. Specifically, the trailing edge split is formed by multiple adjacent small protrusions, wherein each pressure surface trailing edge point 51 is on the inner side of the small protrusion, and the surface where the pressure surface trailing edge point 51 is located is the pressure surface, and the suction surface trailing edge point 52 is on the outer side of the small protrusion, and the surface where the suction surface trailing edge point 52 is located is the suction surface.

[0032] like Figure 1 As shown, in one embodiment, air film holes are provided on the blade 11, the first edge plate 12 and the second edge plate 13. The high-pressure turbine blade 11 adopts a complex cooling form, and air film holes are opened on the blade body and the first edge plate 12 and the second edge plate 13 to ensure a better cooling effect to adapt to high-temperature combustion gas without being ablated. The diameter of the air film hole is 0.4-0.6mm; the trailing edge adopts a split seam design, and the cold air flows out from the trailing edge.

[0033] like Figure 3 As shown, in one embodiment, the number of blades 11 in each blade assembly is 4 to 8, and optionally, the blade assembly is a fan-shaped structure. The number of blades 11 of the present invention is 4, so that the blades 11 are conveniently processed and assembled in a group, thereby reducing the difficulty of processing the entire structure and reducing the leakage loss between the edge plates. At the same time, the process feasibility is improved and the manufacturing cost is reduced. In addition, a sealing groove 110 is provided at the connection between each blade assembly, and the sealing groove 110 is respectively arranged on the first edge plate 12 and the second edge plate 13, and the sealing sheet installed on the sealing groove 110 is used to seal the connection between the blade assemblies. Optionally, it also includes a casing plate 2, the edge of the second edge plate 13 is connected to the casing plate 2, and the groove on the second edge plate 13 cooperates with the boss 16 on the casing plate 2 for circumferential positioning.

[0034] In summary, the turbine guide vane of the present invention is connected to the combustion chamber, and the combustion chamber is located at the front end of the turbine guide vane, and is matched with the combustion chamber through the turbine guide vane, and the blades 11 and the first edge plates 12 and the second edge plates 13 connected to the blades 11 are finely designed, so that under the conditions of high load and high flow in the channel between the blades 11, the load on the blades 11 is evenly distributed, and the aerodynamic parameters on the surface of the blades 11 are reasonably distributed, which reduces the secondary flow loss and the leakage loss between the edge plates, and improves the blade turbine efficiency; the cooling effect is improved, the casting processing technology of the blades 11 is good, and the manufacturing cost is low.

[0035] Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent substitutions for some of the technical features therein; and these modifications or substitutions do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A turbine guide vane, characterized in that: include: A plurality of connected blade assemblies, each of the blade assemblies comprising a plurality of blades (11), a first edge plate (12) and a second edge plate (13), one end of the plurality of blades (11) being connected to the first edge plate (12) and the other end being connected to the second edge plate (13), the first edge plate (12) and the second edge plate (13) being arranged coaxially, the axis on the surface formed by the connection between each blade (11) and the second edge plate (13) being parallel to the axial line of the engine, the angle formed between the tangent on the first edge plate (12) and the radial height of the blade (11) gradually decreasing until the angle gradually increases after the blade (11) is in a preset position, so that the channel area formed between adjacent blades (11) decreases from large to small until the channel area increases from small to large after the blade (11) is in the preset position.

2. The turbine guide vane according to claim 1, characterized in that: The inlet configuration angle of the plurality of blades (11) is 110°-130°.

3. The turbine guide vane according to claim 1, characterized in that: The outlet angles of the plurality of blades (11) gradually decrease.

4. The turbine guide vane according to claim 3, characterized in that: The outlet angle of the plurality of blades (11) is 20°-30°.

5. The turbine guide vane according to claim 1, characterized in that: The installation angle of the blade (11) is 40°-50°.

6. The turbine guide vane according to any one of claims 1 to 5, characterized in that: The thickness of the front half of the cross section (5) of the blade (11) remains constant along the chord length direction or varies within a preset thickness range along the chord length direction.

7. The turbine guide vane according to claim 6, characterized in that: Each blade (11) is installed at a different angle between the first edge plate (12) and the second edge plate (13), so that a plurality of flow channels with different areas are formed between the blades (11).

8. The turbine guide vane according to claim 1, characterized in that: A trailing edge split (50) is provided at the exhaust end of the blade (11), one end of the trailing edge split (50) has a pressure surface trailing edge point (51), and the other end has a suction surface trailing edge point (52), the pressure surface trailing edge point (51) and the suction surface trailing edge point (52) are connected by ribs forming the trailing edge split (50).

9. The turbine guide vane according to claim 1, characterized in that: Air film holes are provided on the blade (11), the first edge plate (12) and the second edge plate (13).

10. The turbine guide vane according to claim 1, characterized in that: The number of blades (11) in each blade assembly is 4 to 8.

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