A transonic turbine rotor blade, turbine rotor and turbine

Abstract

The present invention discloses a transonic turbine rotor blade, a turbine rotor and a turbine. The airfoil of the rotor blade includes a blade leading edge, a blade trailing edge, a blade suction surface and a blade pressure surface. The gap between the blade pressure surface and the blade suction surface of an adjacent rotor blade forms a cascade passage. The cascade passage has a cascade throat, and the cascade throat has a first end on the side of the blade pressure surface and a second end on the side of the blade suction surface. The first end is located at the blade trailing edge, and the second end is located at a position 1 / 6 to 1 / 5 of the blade chord length away from the blade leading edge; the included angle formed by the tangent line of the second end and the tangent line of the blade trailing edge on the same side is 7° to 16°; the included angle formed by the tangent lines of the blade suction surface and the blade pressure surface at the blade trailing edge is 3° to 10°. The present invention can effectively adapt to the flow under strong shock waves in transonic flow, and effectively reduces the flow loss.

Description

Technical Field

[0001] The present invention belongs to the field of engines, and particularly relates to a transonic turbine rotor blade, a turbine rotor and a turbine. Background Art

[0002] An aeroengine can be structurally divided into three major components: a compressor, a combustor, and a turbine. The air flow is pressurized in the compressor and then enters the combustor for combustion, and finally enters the turbine for expansion work to generate the power required by the aircraft. As one of the three major components, the main function of the turbine is to extract the energy from the fluid working medium and output mechanical work in a rotational manner, which is applied to gas turbine engines and aircraft auxiliary power units.

[0003] One of the current major development trends of gas turbine engines is to gradually develop towards higher power-to-weight ratios, higher thermal cycle parameters, and more compact structures. In order to increase the output power of the engine and reduce the engine weight, the temperature before the turbine and the expansion ratio of the turbine are continuously increased. Due to the temperature resistance limitation of the blade material, the temperature before the turbine is usually lower than a certain limit value and cannot be further increased. Therefore, in order to increase the output power of the entire engine, only by increasing the expansion ratio of the turbine, that is, increasing the load of a single row of blades. However, excessive increase in the turbine expansion ratio will cause a large change in the flow state of the entire turbine cascade channel. Especially for a single-stage axial flow turbine with an expansion ratio exceeding 2.5, both the turbine nozzle guide vane and the rotor flow channel may be in a transonic or even supersonic flow state, resulting in a very strong shock wave in the turbine cascade channel, greatly increasing the overall flow loss of the turbine; at the same time, the increase in blade load will increase the leakage flow loss in the rotor tip clearance, resulting in a decrease in turbine performance; on the other hand, there is a shock wave loss in the rotor cascade channel, which makes the rotor outlet parameters change violently, increases the intensity of the secondary flow in the cascade channel, and leads to a relatively large rotor secondary flow loss, affecting the improvement of turbine efficiency.

[0004] In addition, due to the large load of the turbine, the rotor blades are often designed to be relatively solid to improve the work capacity of a single blade. However, this will also bring other problems, such as the weight of a single blade being too heavy, and under high-temperature and high-speed rotation, the stress generated exceeds the bearing range of the blade root tenon, causing great risks. Adjusting the overall center of gravity position of the blade is also a difficult task. Therefore, generally speaking, the design of rotor blades in transonic turbines is more difficult and various factors need to be comprehensively considered.

[0005] In the existing technical solutions, the single-stage expansion ratio of the turbine is usually not high. The turbine rotor blades usually adopt a simple straight solid blade design, and the flow inside the turbine rotor is subsonic flow, with a low ability to accelerate the airflow, which affects the work capacity of the turbine, thereby affecting the improvement of the engine power-to-weight ratio and the scope of use. Secondly, in the existing technical solutions, the design of the turbine rotor blades is simple, the ability to control the secondary flow is weak, and the flow loss is large. Especially, the adaptability to the shock waves existing in the transonic flow is poor, which greatly affects the performance of the turbine. Summary of the Invention

[0006] In view of the above problems, the technical solution adopted by the present invention is: a transonic turbine rotor blade, and the blade profile of the rotor blade includes a blade leading edge, a blade trailing edge, a blade suction surface, and a blade pressure surface.

[0007] The gap between the blade pressure surface and the blade suction surface of the adjacent rotor blades forms a cascade channel. The cascade channel has a cascade throat. The cascade throat has a first end on the side of the blade suction surface and a second end on the side of the blade pressure surface. The first end is located at the blade trailing edge, and the second end is located at a position 1 / 6 to 1 / 5 of the blade chord length away from the blade leading edge.

[0008] The included angle formed by the tangent line of the second end and the tangent line of the blade trailing edge on the same side is 7° to 16°.

[0009] The included angle formed by the tangent lines of the blade pressure surface and the blade suction surface at the blade trailing edge is 3° to 10°.

[0010] Optionally, the second end is located at a position 1 / 5 of the blade chord length away from the blade leading edge.

[0011] Optionally, the cross-sectional area of the blade profile of the rotor blade gradually decreases from the blade root to the blade tip, and the reduction amplitude from the blade root to the blade mid-span is greater than the reduction amplitude from the blade mid-span to the blade tip.

[0012] Optionally, the radius of the blade leading edge gradually decreases from the blade root to the blade tip of the rotor blade, and the reduction amplitude from the blade root to 0.8 times the blade height is greater than the reduction amplitude from 0.8 times the blade height to the blade tip.

[0013] Optionally, the ratio of the cross-sectional areas of the blade profiles at the blade root, blade mid-span, and blade tip of the rotor blade is 2.82:1.54:1.

[0014] Optionally, the ratio of the radii of the blade leading edge at the blade root, blade mid-span, and blade tip is 2.247:1.443:1.

[0015] Optionally, the ratio of the maximum thickness of the blade profile of the rotor blade to the chord length of the rotor blade is 0.1 to 0.3.

[0016] Optionally, the ratios of the maximum thickness of the blade profile of the rotor blade to the blade chord length are 0.295, 0.175, and 0.11 at the blade root, blade middle, and blade tip, respectively.

[0017] Optionally, a weight reduction groove is provided on the rotor blade. The opening side of the weight reduction groove is located at the blade tip end of the rotor blade, and the depth of the weight reduction groove is 1 / 3 - 2 / 5 of the height of the rotor blade.

[0018] Optionally, the blade pressure side and suction side profiles of the weight reduction groove are formed by offsetting the blade profiles of the same height of the rotor blade, and the offset distances are different for different cross - section heights.

[0019] Optionally, the cross - sectional area of the top of the weight reduction groove is 27% of the cross - sectional area of the blade tip of the rotor blade; the cross - sectional area of the bottom of the weight reduction groove is 25% of the cross - sectional area of the cross - section of the rotor blade at the same height.

[0020] Optionally, the wall thicknesses of the blade pressure side and suction side at the top of the weight reduction groove are 25% and 20.8% of the maximum thickness of the blade profile at the blade tip of the rotor blade, respectively; the wall thicknesses of the blade pressure side and suction side at the bottom of the weight reduction groove are 23% and 18.5% of the maximum thickness of the blade profile at the same cross - section height of the rotor blade, respectively.

[0021] Optionally, the cross - sections at the top and bottom of the weight reduction groove are basic cross - sections, and the overall shape of the weight reduction groove is formed by linear interpolation of the two basic cross - sections.

[0022] And, a transonic turbine rotor, the turbine rotor includes a rotor outer ring, a rotor inner ring, and a plurality of the above - mentioned rotor blades. The rotor outer ring and the rotor outer ring are symmetrically arranged around the engine axis. An annular air flow passage is formed between the rotor outer ring and the rotor inner ring. The blade roots of the plurality of rotor blades are fixedly arranged near the rotor inner ring and are uniformly distributed in the circumferential direction between the radial annular passage formed by the rotor inner ring and the rotor outer ring.

[0023] And, a transonic turbine, the turbine includes the above - mentioned turbine rotor.

[0024] Due to the adoption of the above - mentioned technical solutions, the present invention has the following beneficial effects:

[0025] 1. It can effectively adapt to the flow under strong shock waves in transonic flow, and effectively reduce the flow loss;

[0026] 2. It has a reasonable law of change of the root - tip area, and the maximum circular thickness and leading - edge radius of the blade change reasonably from the blade root to the blade tip. The blade is relatively solid as a whole, and the blade has good aerodynamic performance and strength performance, improving the blade life and the reliability of the engine;

[0027] 3. The weight-reducing groove design of the rotor blade further reduces the weight of the blade, improves the blade life, increases the reliability of the engine, and the weight-reducing groove can reduce the leakage flow in the tip clearance and improve the performance of the turbine.

[0028] Other features and advantages of the present invention will be set forth in the following description, and in part will be obvious from the description, or may be learned by practice of the present invention. The objectives and other advantages of the present invention may be realized and attained by the structure particularly pointed out in the specification, claims as well as the drawings. Brief Description of the Drawings

[0029] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can be obtained based on these drawings.

[0030] Figure 1 Shows a three-dimensional structural schematic diagram of a transonic turbine rotor blade according to an embodiment of the present invention;

[0031] Figure 2 Shows a schematic diagram of the blade profile and cascade passage of a rotor blade according to an embodiment of the present invention;

[0032] Figure 3 Shows a diagram of the airflow in the cascade passage of a rotor blade according to an embodiment of the present invention;

[0033] Figure 4 Shows a schematic diagram of the blade profiles at the root, middle, and tip cross-sections of a rotor blade according to an embodiment of the present invention;

[0034] Figure 5 Shows a diagram of the variation law of the root, middle, and tip surface area of a rotor blade according to an embodiment of the present invention;

[0035] Figure 6 Shows a diagram of the variation law of the leading edge radius of a rotor blade at the root, middle, and tip according to an embodiment of the present invention;

[0036] Figure 7 Shows a diagram of the variation law of the relative thickness of a rotor blade at the root, middle, and tip according to an embodiment of the present invention;

[0037] Figure 8 Shows a perspective view of a rotor blade according to an embodiment of the present invention;

[0038] Figure 9 Shows a schematic diagram of the profile line of the weight-reducing groove according to an embodiment of the present invention;

[0039] Figure 10 Shows a two-dimensional schematic diagram of a turbine flow passage according to an embodiment of the present invention;

[0040] Figure 11 Shows the transonic turbine test efficiency diagram according to an embodiment of the present invention.

[0041] Wherein, 1. Leading edge of the blade, 2. Trailing edge of the blade, 3. Suction side of the blade, 4. Pressure side of the blade, 5. Cascade passage, 6. Throat of the cascade, 7. Leading edge of the weight reduction groove, 8. Trailing edge of the weight reduction groove, 9. Bottom of the weight reduction groove, 10. Top of the weight reduction groove, 11. Depth of the weight reduction groove, 12. Height of the rotor blade, 13. Suction side of the weight reduction groove, 14. Pressure side of the weight reduction groove, 15. Wall thickness of the suction side of the weight reduction groove, 16. Wall thickness of the pressure side of the weight reduction groove, 17. Maximum thickness of the airfoil. Detailed implementation manners

[0042] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0043] As Figure 1 Shown in the three-dimensional structural schematic diagram of the transonic turbine rotor blade, the airfoil of the transonic turbine rotor blade in the embodiment of the present invention includes the leading edge 1 of the blade, the trailing edge 2 of the blade, the suction side 3 of the blade, and the pressure side 4 of the blade.

[0044] As Figure 2 Shown in the schematic diagram of the rotor blade airfoil and the cascade passage, the gap between the pressure side 4 of the rotor blade and the suction side of the adjacent rotor blade forms the cascade passage 5. In this embodiment, the rotor blade adopts a convergent airfoil design. The cascade passage 5 has a throat 6 of the cascade. The throat of the cascade has a first end on the suction side of the blade and a second end on the pressure side of the blade. The first end is located at the trailing edge 2 of the blade. Therefore, along the air flow direction, the area of the air flow passage is always gradually decreasing, that is, there is Figure 2 The relationship between a1 and a2 in is a1 > a2. After the air flow enters the cascade passage and before reaching the throat of the cascade, the air flow velocity is continuously increasing and reaches the critical Mach number at the throat of the cascade. In order to enable the air flow to continue to accelerate after flowing through the throat of the cascade, the principle similar to that of a Laval nozzle is adopted, and the second end of the throat of the cascade is arranged at a distance of 1 / 6 - 1 / 5 of the blade chord length from the leading edge 1 of the blade. In this embodiment, the second end is located at a distance of 1 / 5 of the blade chord length from the leading edge 1 of the blade, that is Figure 2As shown by point M in the figure, b1 = 1 / 5b. In this way, when the fluid flows out of the throat of the blade, the blade back profile expands outward after flowing through the throat, and the blade basin is free from blade profile constraints. The flow area will suddenly increase and can continue to expand and accelerate at the oblique cut position, as shown in the following example: Figure 3 The diagram of the airflow within the rotor blade cascade passages is shown. This ultimately allows the airflow to reach a Mach number of approximately 1.05 upon entering the downstream rotor blade leading edge. A high rotor exit Mach number indicates sufficient work is being performed within the rotor, improving the airflow's ability to perform work and increasing the turbine's output. Therefore, designing the throat position at this location can increase turbine output.

[0045] Since the flow inside the turbine blade channel is transonic, in order to reduce the blade flow loss at high Mach numbers, such as Figure 2 As shown, the curvature of the blade tip of this rotor blade design changes minimally after the throat point M, resulting in a nearly straight line. The degree of blade back curvature is represented by the trailing edge bend angle δ, defined as the angle formed by the tangent line through point M and the tangent line on the same side of the blade's trailing edge. δ ranges from 7° to 16°, with larger values indicating greater back curvature. In this embodiment, δ = 9°.

[0046] Secondly, due to the transonic flow in the turbine rotor blade channel, there are relatively obvious shock waves. In order to reduce the losses caused by shock waves and wakes, the trailing edge wedge angle α is defined as the angle between the tangents of the blade back profile and the blade basin profile at the small circle of the blade trailing edge. Using a smaller blade trailing edge wedge angle α can effectively reduce wake losses. If the trailing edge wedge angle α is too large, there will be a large collision angle between the airflows flowing out of the blade basin and the blade back, and a large recirculation area will be formed at the intersection, which will greatly increase the flow losses. Usually in transonic blade design, the trailing edge wedge angle α is between 3° and 10°. In this embodiment, α = 6°.

[0047] Since the flow inside the turbine rotor is transonic, the tip tangential velocity of the rotor blade is high and the stress at the rotor root is large. In order to ensure that the rotor blade still has good strength performance under high speed conditions, such as Figure 4 The blade profile schematic diagram of the root, middle and tip sections of the rotor blade is shown, and the cross-sectional area of the rotor blade gradually decreases from the root to the tip. When the rotor blade rotates, the linear velocity of the blade tip along the circumferential tangent is very high, and the tangential velocity of the blade tip is more than 1.3 times the tangential velocity of the blade root. Therefore, the blade profile adopts a design method with a large root and a small tip, which can increase the fracture resistance of the entire rotor blade and extend the life of the rotor blade. Usually, the area ratio of the root-tip section should be above 1.7. In this embodiment, the profile area ratio of the root-tip blade is 2.82, which makes the entire rotor blade have good strength performance. And the deceleration amplitude from the root to the middle of the blade is greater than the deceleration amplitude from the middle of the blade to the tip, such as Figure 5The figure showing the variation law of the areas of the root, middle, and tip profiles of the rotor blade. The profile area decreases rapidly from the root to 50% of the blade height, and the rate of decrease in the profile area slows down from 50% of the blade height to the tip. This design can endow the rotor blade with good strength performance while taking into account the aerodynamic performance under transonic flow. In this embodiment, the ratio of the profile cross-sectional areas at the root, middle, and tip of the rotor blade is 2.82:1.54:1.

[0048] Furthermore, when designing the radius of the blade leading edge, the strength and aerodynamic performance of the blade are also taken into account. The radius of the blade leading edge gradually decreases from the root to the tip of the rotor blade, and the reduction amplitude from the root to 0.8 times the blade height is greater than that from 0.8 times the blade height to the tip. As Figure 6 shown in the figure of the variation law of the radius of the rotor blade leading edge at the root, middle, and tip. In the region from the root to 0.8 times the blade height, the radius of the leading edge basically shows a linear change, and in the region from 0.8 times the blade height to the tip, the change amplitude of the leading edge radius decreases. In this embodiment, the ratio of the radii of the blade leading edge at the root, middle, and tip is 2.247:1.443:1. Adopting this design can optimize the comprehensive performance of the rotor blade.

[0049] Define that a perpendicular line is drawn from any point on the suction side of the blade to the pressure side of the blade, and the length of the perpendicular line segment connecting the suction side and the pressure side of the blade is defined as the thickness of the rotor blade profile. As Figure 2 shown, the ratio of the maximum thickness Cmax of the profile to the chord length b of the blade is defined as the relative thickness. The smaller the relative thickness value, the more slender the profile, and the larger the relative thickness value, the fatter the profile. Generally, the relative thickness Cmax / b of the rotor blade profile is 0.1 - 0.3. As Figure 7 shown in the figure of the variation law of the relative thickness of the rotor blade at the root, middle, and tip. In this embodiment, the ratios of the maximum thickness of the rotor blade profile to the chord length b at the root, middle, and tip are 0.295, 0.175, and 0.11 respectively. The relative thickness of the tip section is the smallest, and the airflow in the tip region is supersonic. A slender profile is more conducive to reducing profile losses. Similar to the root tip area ratio, the thicknesses of the leading and trailing edges of the blade also gradually decrease along the blade height direction. In order to reduce the wake loss when the airflow flows out of the blade, the smaller the trailing edge radius R2, the better, but limited by the processing level, if the trailing edge radius is too small, it will be impossible to process. Therefore, the trailing edge radius R2 is usually above 0.2 mm. In this example, in order to balance the processing level, the trailing edge radius is taken as 0.35 mm for all. Based on the ratio of the radii of the blade leading edge at the root, middle, and tip being 2.247:1.443:1, the leading edge radii of the blade in this embodiment are 1.09 mm, 0.7 mm, and 0.485 mm respectively.

[0050] As Figure 8Perspective view of the rotor blade shown, with weight-reducing grooves provided on the rotor blade. The opening side of the weight-reducing grooves is located at the tip end of the rotor blade, further improving the strength performance of the turbine rotor blade, increasing the blade life, and controlling the tip leakage flow. The depth 11 of the weight-reducing grooves is 1 / 3 to 2 / 5 of the height 12 of the rotor blade. In this embodiment, the depth 11 of the weight-reducing grooves is 33.5% of the height of the rotor blade.

[0051] Secondly, Figure 9 The schematic diagram of the weight-reducing groove profile is shown. Combining Figure 8 and Figure 9 As shown, the profiles of the suction side 13 and pressure side 14 of the weight-reducing grooves are formed by offsetting the blade profiles at the same height of the rotor blade, and the offset distances at different cross-sectional heights are different. The cross-sectional area of the top 10 of the weight-reducing groove is 27% of the tip cross-sectional area of the rotor blade; the cross-sectional area of the bottom 9 of the weight-reducing groove is 25% of the cross-sectional area of the rotor blade at the same height. At the top 10 of the weight-reducing groove, the wall thickness 15 of the suction side of the weight-reducing groove and the wall thickness 16 of the pressure side of the weight-reducing groove are 25% and 20.8% respectively of the maximum thickness 17 of the blade profile at the tip of the rotor blade; at the bottom 9 of the weight-reducing groove, the wall thickness 15 of the suction side of the weight-reducing groove and the wall thickness 16 of the pressure side of the weight-reducing groove are 23% and 18.5% respectively of the maximum thickness 17 of the blade profile at the same cross-sectional height of the rotor blade. At the top 10 of the weight-reducing groove, the diameter of the leading edge 7 of the weight-reducing groove is 55.6% of the diameter of the leading edge 1 of the blade at the same cross-sectional height, and the diameter of the trailing edge 8 of the weight-reducing groove is 50% of the diameter of the trailing edge 2 of the blade at the same cross-sectional height; at the bottom 9 of the weight-reducing groove, the diameter of the leading edge 7 of the weight-reducing groove is 29% of the diameter of the leading edge 1 of the blade at the same cross-sectional height, and the diameter of the trailing edge 8 of the weight-reducing groove is 55.6% of the diameter of the trailing edge 2 of the blade at the same cross-sectional height. The cross-sections at the top and bottom of the weight-reducing groove are the basic cross-sections, and the overall shape of the weight-reducing groove is formed by linear interpolation of these two basic cross-sections. The weight-reducing groove designed by this method can reduce the weight of the rotor blade, meet the reliability requirements of the engine, reduce the tip leakage loss, and improve the efficiency of the rotor blade.

[0052] Furthermore, the transonic turbine rotor of the embodiment of the present invention, as Figure 10 shown in the two-dimensional schematic diagram of the turbine flow passage, the turbine rotor includes a rotor outer ring, a rotor inner ring, and a plurality of the transonic turbine rotor blades as described in the above embodiments. The rotor outer ring and the rotor outer ring are symmetrically arranged around the engine axis. An annular air flow passage is formed between the rotor outer ring and the rotor inner ring. The blade roots of the plurality of rotor blades are fixedly arranged close to the rotor inner ring and are evenly distributed in the radial annular passage formed between the rotor inner ring and the rotor outer ring along the circumferential direction.

[0053] Still further, the transonic turbine of the embodiment of the present invention includes the transonic turbine rotor as described in the above embodiments.

[0054] Implementation effect of this embodiment:

[0055] The aerodynamic performance of a turbine blade is generally evaluated by the total pressure loss coefficient or the total turbine efficiency. The calculation formula for the total pressure loss coefficient of the blade is:

[0056] ξ = 1 - P t2 / P t1

[0057] Where, P t2 is the total pressure at the blade outlet, P t1 is the total pressure at the blade inlet, and ξ is the total pressure loss when the air flow passes through the blade, reflecting the aerodynamic performance of the blade. The larger ξ is, the worse the performance of the blade.

[0058] The following Table 1 gives a comparison of the calculation results of the original conventional design and the design of the present invention. It can be seen from the table that by using the blade design method of the present invention, the total pressure loss has been greatly reduced, approximately by about 7% under the design point conditions.

[0059] Table 1 Total pressure loss of the blade

[0060] Operating condition 1 (design point) Operating condition 2 Operating condition 3 Original conventional design 0.0644 0.0742 0.0855 Design of the present invention 0.060 0.0695 0.081

[0061] In addition, a performance test was carried out on the turbine component containing the blade designed by the present invention. The test results of the component performance show that the performance of the turbine component containing the blade of the present invention reaches the designed index value. The calculated value of the efficiency of the turbine component of the present invention is 0.879. As Figure 11 shown in the transonic turbine test efficiency diagram, the test efficiency value of the present invention is 0.8796, and the required value of the design index is 0.875. The calculated value of the efficiency of the turbine component using the original conventional design is 0.873. It can be seen that the blade design method of the present invention can effectively improve the turbine performance, and both the aerodynamic performance of the single-row blade and the overall performance of the turbine component have been improved.

[0062] Where, the calculated value of the turbine component efficiency η t uses the following formula:

[0063]

[0064] In the formula, Ne: the total power measured in the test; T t0 : the total temperature of the inlet air flow; π t : the total pressure expansion ratio, = P t0 / P t4 ; η m : the mechanical efficiency; W: the intake air flow of the model turbine; k: the adiabatic index 1.378; R: the gas constant 287.320 J / (kg·K).

[0065] In terms of strength performance, by adopting design methods such as the appropriate root tip area ratio, maximum circle distribution method, and weight reduction design designed in the present invention, the aerodynamic performance of the blade is greatly improved, and the blade strength performance is also improved significantly. The blade weight is reduced by 19.8% compared with the original conventional design. The following Table 2 shows the strength calculation results designed in the present invention. As shown in Table 2, the stress values designed in the present invention are greatly reduced. The maximum equivalent stress is reduced by 183 MPa, the maximum radial stress is reduced by 295 MPa, and the maximum principal stress is reduced by 294 MPa. The reduction of stress values is very beneficial to the life and safety of the blade. Therefore, it can be seen that the design of the rotor blade of the present invention can greatly improve the strength performance.

[0066] Comparison of Stress Calculation Results in Table 2

[0067]

[0068] In summary, by adopting the transonic turbine rotor mentioned above, the aerodynamic performance and strength performance of the rotor blade of the present invention are greatly improved, and the comprehensive performance of the entire turbine engine is greatly improved.

[0069] It should be noted that the structures, proportions, sizes, etc. shown in the drawings of this specification are only used to cooperate with the content disclosed in the specification for those who are familiar with this technology to understand and read, and are not used to limit the limiting conditions under which the present invention can be implemented. Therefore, they do not have technical essential significance. Any modification of the structure, change of the proportional relationship, or adjustment of the size, without affecting the effects that the present invention can produce and the purposes that can be achieved, should still fall within the scope covered by the technical content disclosed in the present invention. At the same time, the terms such as "upper", "lower", "left", "right", "middle", and "one" cited in this specification are only for the convenience of clear narration, and are not used to limit the scope under which the present invention can be implemented. The change or adjustment of their relative relationships, without substantial change in the technical content, should also be regarded as the scope under which the present invention can be implemented.

[0070] The above description is only a preferred embodiment of the present invention, and does not make any form of limitation to the present invention. Although the present invention has been disclosed above with a preferred embodiment, it is not used to limit the present invention. Any person skilled in the art, without departing from the scope of the technical solution of the present invention, can make some changes or modifications to the above-disclosed technical content to be equivalent change equivalent embodiments. However, as long as it does not depart from the content of the technical solution of the present invention, any simple modification, equivalent change, and modification made to the above embodiments according to the technical essence of the present invention still fall within the scope of the technical solution of the present invention.

Claims

1. A transonic turbine rotor blade, the airfoil of the rotor blade including a blade leading edge, a blade trailing edge, a blade suction surface, and a blade pressure surface, characterized in that a gap between the blade pressure surface and the blade suction surface of an adjacent rotor blade forms a cascade passage, the cascade passage having a cascade throat, the cascade throat having a first end on the side of the blade pressure surface and a second end on the side of the blade suction surface, the first end being located at the blade trailing edge, and the second end being located at a position 1 / 6 to 1 / 5 of the blade chord length away from the blade leading edge; an included angle formed by a tangent line of the second end and a tangent line of the blade trailing edge on the same side is 7° to 9°; an included angle formed by a tangent line of the blade pressure surface and a tangent line of the blade suction surface at the blade trailing edge is 3° to 6°; the airfoil cross-sectional area of the rotor blade gradually decreases from the blade root to the blade tip, and the reduction amplitude from the blade root to the blade mid-span is greater than that from the blade mid-span to the blade tip; the radius of the blade leading edge gradually decreases from the blade root to the blade tip of the rotor blade, and the reduction amplitude from the blade root to 0.8 times the blade height is greater than that from 0.8 times the blade height to the blade tip.

2. The transonic turbine rotor blade according to claim 1, characterized in that, The second end is located at a position 1 / 5 of the blade chord length away from the blade leading edge.

3. The transonic turbine rotor blade according to claim 1, characterized in that, The ratio of the airfoil cross-sectional areas at the blade root, blade mid-span, and blade tip of the rotor blade is 2.82:1.54:

1.

4. The transonic turbine rotor blade according to claim 1, wherein, The ratio of the radii of the blade leading edge at the blade root, blade mid-span, and blade tip is 2.247:1.443:

1.

5. The transonic turbine rotor blade according to claim 1, characterized in that, The ratio of the maximum thickness of the airfoil of the rotor blade to the chord length of the rotor blade is 0.1 to 0.

3.

6. The transonic turbine rotor blade according to claim 5, wherein, The ratios of the maximum thickness of the airfoil to the chord length at the blade root, blade mid-span, and blade tip of the rotor blade are 0.295, 0.175, and 0.11 respectively.

7. The transonic turbine rotor blade according to any one of claims 1 to 6, characterized in that, A weight-reducing groove is provided on the rotor blade, the opening side of the weight-reducing groove being located at one end of the blade tip of the rotor blade, and the depth of the weight-reducing groove being 1 / 3 to 2 / 5 of the height of the rotor blade.

8. The transonic turbine rotor blade according to claim 7, characterized in that, The airfoil lines of the pressure surface and suction surface of the weight-reducing groove are formed by offsetting the airfoils of the cross-sections at the same height of the rotor blade, and the offset distances at different cross-sectional heights are different.

9. The transonic turbine rotor blade according to claim 7, wherein The cross-sectional area of the top of the weight-reducing groove is 27% of the cross-sectional area of the blade tip of the rotor blade; the cross-sectional area of the bottom of the weight-reducing groove is 25% of the cross-sectional area of the cross-section at the same height of the rotor blade.

10. The transonic turbine rotor blade according to claim 8, characterized in that, The wall thicknesses of the pressure surface and suction surface at the top of the weight-reducing groove are 25% and 20.8% respectively of the maximum thickness of the airfoil at the blade tip of the rotor blade; the wall thicknesses of the pressure surface and suction surface at the bottom of the weight-reducing groove are 23% and 18.5% respectively of the maximum thickness of the airfoil at the same cross-sectional height of the rotor blade.

11. The transonic turbine rotor blade according to any one of claims 8 to 10, characterized in that, The cross-sections at the top and bottom of the weight-reducing groove are basic cross-sections, and the overall shape of the weight-reducing groove is formed by linear interpolation of the two basic cross-sections.

12. A transonic turbine rotor, characterized in that, The turbine rotor includes a rotor outer ring, a rotor inner ring, and a plurality of rotor blades as described in any one of claims 1 to 11. The rotor outer ring and the rotor outer ring are symmetrically arranged around the engine axis. An annular air flow passage is formed between the rotor outer ring and the rotor inner ring. The blade roots of the plurality of rotor blades are fixedly arranged close to the rotor inner ring and are uniformly distributed in the circumferential direction between the radial annular passage formed by the rotor inner ring and the rotor outer ring.

13. A transonic turbine, characterized in that, The turbine includes a turbine rotor as described in claim 12.

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