Engine turbine with high-altitude performance and its design method

Abstract

The present invention belongs to the technical field of aeroengines, and particularly relates to an engine turbine with high-altitude performance and a design method thereof. The engine turbine includes a first-stage turbine, a second-stage turbine, an internal flow passage, and an external flow passage. The first-stage turbine and the second-stage turbine are respectively evenly distributed along the radial direction of the annular passage. The first-stage turbine includes a first-stage rotor and a first-stage stator vane. The second-stage turbine includes a second-stage rotor and a second-stage stator vane. The first-stage stator vane, the first-stage rotor, the second-stage stator vane, and the second-stage rotor are arranged in sequence from the side close to the combustion chamber to the side far from the combustion chamber. The turbine blade profile of the present invention can effectively adapt to the flow under strong transonic shock waves, effectively reduce the flow loss, reduce the exhaust loss, and improve the turbine performance. In addition, the expansion ratio and load coefficient of the turbine of the present invention are more reasonably selected, which can take into account the performance of the turbine during high-altitude operation, improve the adaptability of the last stage of the turbine to the sharp increase in high-altitude expansion ratio and poor conditions such as low Reynolds number at high altitude, and greatly improve the turbine performance.

Description

Technical Field

[0001] The present invention belongs to the technical field of aeroengines, and particularly relates to an engine turbine with high-altitude performance and a design method thereof. 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 used in gas turbine engines and aircraft auxiliary power units. Currently, one major development trend of gas turbine engines is to gradually develop towards higher power-to-weight ratios, higher thermal cycle parameters, and more compact structures, resulting in increasingly harsh and severe working conditions for turbine components. Moreover, with the increasingly urgent demand for all-weather multi-purpose aircraft, the operating environment of the aircraft will become more and more complex, its operating envelope will gradually expand, and there are higher requirements for the operating range of the engine. Therefore, in the turbine design process, not only the increasing turbine load needs to be considered, but also the working conditions in various complex environments need to be taken into account, such as the use in high-altitude environments.

[0003] In order to increase the output power of the engine and reduce the weight of the engine, 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, 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 great changes in the flow state of the entire cascade passage of the turbine. Especially for an axial flow turbine with a double-stage turbine expansion ratio exceeding 7, both the turbine nozzle guide vane and the rotor flow passage are in a transonic or even supersonic flow state, resulting in strong shock waves in the turbine cascade passage, greatly increasing the overall flow loss of the turbine. If the load between the turbine stages cannot be effectively and well distributed, it will surely lead to poor performance of the turbine. At the same time, the exhaust velocity at the outlet of the transonic turbine is high, resulting in large exhaust losses, which will affect the overall performance of the turbine.

[0004] On the other hand, the flight altitude of current aircraft is getting higher and higher, especially for civil airliners, large transport aircraft, bombers, etc., whose operating envelopes are getting wider and the operating altitude can reach 20 km. In the existing technical solutions, when the engine operates in a high-altitude environment, the turbine efficiency drops sharply, the working state of the turbine seriously deviates from the design point, and the internal flow loss of the transonic turbine is large and the exhaust velocity is high, affecting the normal output of the turbine power and greatly deteriorating the performance of the turbine. Summary of the Invention

[0005] In view of the above problems, on the one hand, the present invention discloses an engine turbine with high-altitude performance. The engine turbine is arranged in an annular channel at the end of the combustion chamber flame tube of the engine. The annular channel is centered on the engine axis. The side close to the axis is set as the inner flow path, and the side far from the axis is set as the outer flow path. The engine turbine includes a first-stage turbine, a second-stage turbine, an inner flow path, and an outer flow path. The first-stage turbine and the second-stage turbine are evenly distributed along the radial direction of the annular channel. Among them,

[0006] The first-stage turbine includes a first-stage rotor and a first-stage stator vane. The second-stage turbine includes a second-stage rotor and a second-stage stator vane. The first-stage stator vane, the first-stage rotor, the second-stage stator vane, and the second-stage rotor are arranged in sequence from the side close to the flame tube to the side far from the flame tube.

[0007] Furthermore, the rotor blades of the first-stage rotor and the second-stage rotor include a blade root and a blade tip extending along the rotor blade body. The direction of the blade tip rotates clockwise relative to the direction of the blade root. The bending and torsion angle between the blade tip and the blade root is 25-35°, preferably 30°.

[0008] Furthermore, the body of the rotor blade gradually narrows during the process of the blade root of the first-stage rotor and the second-stage rotor extending towards the blade tip.

[0009] Furthermore, the body of the stator vane gradually widens during the process of the blade root of the first-stage stator vane and the second-stage stator vane extending towards the blade tip.

[0010] Furthermore, a blade tip groove is provided on the end face of the blade tip of the rotor blade of the first-stage rotor.

[0011] Furthermore, the intersection line between the section of the first-stage stator vane close to the flame tube and the body of the first-stage stator vane is set as a first leading edge profile line. The included angle A between the first leading edge profile line and the vertical direction is 5-10°.

[0012] Furthermore, the intersection line between the section of the second-stage stator vane close to the flame tube and the body of the second-stage stator vane is set as a second leading edge profile line. The included angle B between the second leading edge profile line and the outer flow path is 90°.

[0013] Still further, the annular channel at the end of the second-stage rotor is set as an expansion type, and the air flow passage area between the outer flow path and the inner flow path on the side far from the flame tube along the air flow direction gradually increases.

[0014] On the other hand, the present invention also proposes a design method for an engine turbine with high-altitude performance. The design method includes:

[0015] Zero-dimensional parameter analysis; determine the load coefficient, flow coefficient, and energy reaction degree of the first-stage turbine and the second-stage turbine respectively according to the design indexes of the engine turbine;

[0016] One-dimensional parameter calculation; according to the basic principles of turbine aerodynamics, obtain the aerodynamic parameters and geometric parameters of the turbine blades of the first-stage turbine and the second-stage turbine of the engine respectively;

[0017] Three-dimensional modeling design; calculate the inlet and outlet aerodynamic parameters and geometric parameters of the first-stage turbine and the second-stage turbine according to the load coefficient, flow coefficient, energy reaction degree, aerodynamic parameters, and geometric parameters, and conduct three-dimensional modeling design of the engine turbine.

[0018] Furthermore, the design indexes of the engine turbine include the flow rate index, expansion ratio index, and adiabatic efficiency index of the first-stage turbine and the second-stage turbine

[0019] The turbine blade profile of the present invention can effectively adapt to the flow under strong transonic shock waves, effectively reduce the flow loss, reduce the exhaust loss, and improve the turbine performance; in addition, the expansion ratio and load coefficient of the turbine of the present invention are more reasonably selected, which can take into account the performance of the turbine during high-altitude operation, improve the adaptability of the last stage of the turbine to the sharp increase in high-altitude expansion ratio and poor conditions such as low Reynolds number at high altitude, and greatly improve the turbine performance.

[0020] Other features and advantages of the present invention will be described in the following specification, and, in part, will be obvious from the specification, or will be understood by implementing the present invention. The objectives and other advantages of the present invention can be achieved and obtained by the structure pointed out in the specification, claims, and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] 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, other drawings can be obtained based on these drawings without creative efforts.

[0022] Figure 1 Shows a schematic diagram of the combustion chamber structure according to an embodiment of the present invention;

[0023] Figure 2 Shows a schematic diagram of the turbine guide vane and rotor blade according to an embodiment of the present invention;

[0024] Figure 3 Shows a schematic diagram of the blade profile of the root, middle, and tip sections of the turbine guide vane according to an embodiment of the present invention;

[0025] Figure 4Shows the schematic diagrams of the blade profiles at the root, middle, and tip sections of a turbine rotor according to an embodiment of the present invention;

[0026] Figure 5 Shows the schematic structural diagram of a first-stage rotor according to an embodiment of the present invention;

[0027] Figure 6 Shows the schematic diagram of the turbine design process according to an embodiment of the present invention;

[0028] Figure 7 Shows the efficiency expansion ratio curve graph of a turbine according to a preferred example of an embodiment of the present invention.

[0029] In the drawings: 1, combustion chamber; 100, flame tube; 101, head of the flame tube; 102, outer ring of the flame tube; 103, inner ring of the flame tube; 104, large elbow; 105, small elbow; 110, annular channel; 111, outer flow channel; 112, inner flow channel; 121, first-stage rotor; 122, second-stage rotor; 12a, root section of the turbine rotor blade; 12b, middle section of the turbine rotor blade; 12c, tip section of the turbine rotor blade; 131, first-stage guide vane; 1311, first leading-edge profile; 132, second-stage guide vane; 1321, second leading-edge profile; 13a, root section of the turbine guide vane blade; 13b, middle section of the turbine guide vane blade; 13c, tip section of the turbine guide vane blade. Detailed implementation manners

[0030] 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.

[0031] The present invention provides a turbine of an engine with high-altitude performance. As an auxiliary power device of an aeroengine, the engine turbine is applied to the end of the flame tube 100 of the combustion chamber 1 of the aeroengine, as Figure 1 shown, which is the schematic structural diagram of the combustion chamber according to the embodiment of the present invention. Among them, the flame tube 100 includes an inner ring 103 of the flame tube, an outer ring 102 of the flame tube, and a head 101 of the flame tube. The inner ring 103 of the flame tube and the outer ring 102 of the flame tube form a flow channel with an annular structure centered on the engine axis. The end of the flow channel is connected to the annular channel 110 through a large elbow 104 and a small elbow 105 in a "C" shape. Inside the flame tube 100, high-temperature gas flows from the flow channel to the annular channel 110, that is, Figure 1 the gas flow direction G in

[0032] In one embodiment of the present invention, reference is made to Figure 1 As shown, the first-stage turbine is connected to the end of the combustion chamber 100. The high-temperature and high-pressure gas flowing out of the combustion chamber 100 enters the first-stage nozzle guide vane 131 for expansion and acceleration. During the process of the gas flowing out of the first-stage nozzle guide vane 131, it can provide a suitable pre-whirl angle of the air flow for the inflow of the first-stage rotor 121. The gas enters the first-stage rotor 121 and impacts the rotor blades to rotate, thereby generating shaft work and driving the power output of the compressor. The functions of the second-stage nozzle guide vane 132 and the second-stage rotor 122 are similar to those of the first-stage nozzle guide vane 131 and the first-stage rotor 121, and will not be elaborated here.

[0033] Specifically, the engine turbine includes a first-stage turbine, a second-stage turbine, an internal flow path 112, and an external flow path 111. The first-stage turbine and the second-stage turbine are evenly distributed along the radial direction of the annular passage 110. In addition, the engine turbine is classified according to the type of blades, including a turbine rotor 12 and a turbine nozzle guide vane 13. Among them, the turbine rotor 12 includes a first-stage rotor 121 and a second-stage rotor 121, and the turbine nozzle guide vane 13 includes a first-stage nozzle guide vane 131 and a second-stage nozzle guide vane 132. The first-stage nozzle guide vane 131, the first-stage rotor 121, the second-stage nozzle guide vane 132, and the second-stage rotor 121 are arranged in sequence from the side close to the combustion chamber 100 to the side far from the combustion chamber 100.

[0034] In one embodiment of the present invention, the rotor blades of the turbine rotor 12 include a blade root and a blade tip extending along the rotor blade body. Specifically, in order to effectively control the flow loss in the turbine internal flow field, as Figure 2 shown, both the first-stage rotor 121 and the second-stage rotor 121 adopt obvious bow-twist characteristics in the blade profile. Referring to Figure 4 shown, it is a schematic diagram of the blade root, blade middle, and blade tip cross-sections of the turbine rotor blade according to an embodiment of the present invention. The turbine rotor blade middle cross-section 12b and the turbine rotor blade tip cross-section 12c of the rotor blade need to be bow-twisted by a certain angle in the clockwise direction on the basis of the turbine rotor blade root cross-section 12a. Preferably, the bow-twist angle between the blade tip and the blade root is 30°. The bow-twist configuration of the rotor blade can effectively control the flow loss inside the annular passage 110 and meet the requirements of transonic flow.

[0035] In one embodiment of the present invention, reference is made to Figure 1 shown, since the flow path curvature at the connection between the first-stage nozzle guide vane 131 and the end of the combustion chamber 100 is too large, it is easy to cause air flow separation. Therefore, during the design process, the leading edge of the first-stage nozzle guide vane 131 is set to be inclined forward, that is, the leading edge profile 1311 of the first-stage nozzle guide vane 131 has an angle A of 5-10° with the vertical direction;

[0036] In addition, as Figure 3As shown, the transverse widths of the root section 13a, the middle section 13b, and the tip section 13c of the turbine guide vane 13 gradually increase, which means that the body of the guide vane gradually widens during the process of the root of the guide vane of the turbine guide vane 13 extending towards the tip. This is used to better adapt to the secondary flow of the end wall and reduce losses.

[0037] In an embodiment of the present invention, referring to Figure 1 As shown, a second leading edge profile 1321 is provided on one side of the secondary guide vane 132 close to the combustion chamber 100. The second leading edge profile 1321 forms an angle B with the external flow passage 111 at the position of the secondary guide vane 132, and the value of the angle B is 90°. This is used to suppress the flow separation at the position of the secondary guide vane 132 and improve the turbine performance. Additionally, referring to Figure 4 As shown, during the process of the root of the turbine rotor 12 extending towards the tip, the body of the rotor blade gradually narrows.

[0038] It should be noted that due to the high-speed rotation of the rotor blade, the tangential speed at the tip of the blade is high, while the tangential speed at the root of the blade is much lower. Considering the strength problem of the blade, it is necessary to reduce the weight at the tip of the blade. Therefore, the blades of the first-stage rotor 121 also adopt the form of a tip groove, as Figure 5 shown, that is, a groove is dug in the top area of the blade. The depth of the groove is 1 / 3 - 2 / 5 of the height of the rotor blade, and in this example, it is 33.5%. The groove is obtained by linear interpolation from two basic sections at the root and the tip; this can not only reduce the weight of the rotor blade but also be beneficial to suppressing the leakage flow in the tip clearance of the rotor blade.

[0039] As an embodiment of the present invention, the first-stage rotor 121 and the second-stage rotor 122 rotate in the same direction and have the same rotational speed. Since the exhaust speed of the transonic turbine is high, in order to reduce the exhaust speed, the annular passage 110 at and after the position of the second-stage rotor 122 is designed to be divergent, that is, the airflow passage area gradually increases, which can achieve the effect of pressure recovery and deceleration to reduce losses.

[0040] In addition, to cope with the influence of high altitude and low Reynolds number on the last stage of the turbine and suppress the separation effect of the flow on the blade surface caused by the low Reynolds number, the last row of blades of the turbine is shaped in a pre-loaded form, that is, the airflow does more work in the first half of the blade. The airflow velocity is relatively low at the leading edge of the blade, and then the airflow accelerates rapidly in the second half. This design can make the airflow transition from laminar flow to turbulent flow in advance at the leading edge of the blade, enhancing the anti-separation ability of the airflow, and thus reducing the negative impact brought by high altitude and low Reynolds number.

[0041] In an embodiment of the present invention, the present invention also proposes a design method for an engine turbine with high-altitude performance, and the method includes:

[0042] Step 1: Zero-dimensional parameter analysis; determine the load coefficient, flow coefficient, and energy reaction degree of the first-stage turbine and the second-stage turbine respectively according to the design indexes of the engine turbine;

[0043] Step 2: One-dimensional parameter calculation; obtain the aerodynamic parameters and geometric parameters of the turbine blades of the first-stage turbine and the second-stage turbine of the engine respectively according to the basic principles of turbine aerodynamics;

[0044] Step 3: Three-dimensional modeling design; calculate the inlet and outlet aerodynamic parameters and geometric parameters of the first-stage turbine and the second-stage turbine according to the load coefficient, flow coefficient, energy reaction degree, aerodynamic parameters, and geometric parameters, and perform three-dimensional modeling design of the engine turbine.

[0045] Furthermore, for a transonic turbine with high-altitude performance,

[0046] Specifically, Step 1 includes: according to the turbine aerodynamic design method, obtain the expansion ratio and efficiency of the first-stage turbine, and then obtain the expansion ratio and efficiency of the second-stage turbine from the total expansion ratio and total efficiency of the turbine. Since the turbine needs to work in a high-altitude environment, the turbine expansion ratio will be about 30% higher than that in the ground state, and the increased expansion ratio will all be added to the second-stage turbine (i.e., the last-stage turbine), resulting in a sharp decline in turbine performance. Therefore, when distributing the expansion ratio, the expansion ratio of the first-stage turbine is large and the expansion ratio of the second-stage turbine is small, and the expansion ratio of the first-stage turbine is about 10%-20% smaller than that of the second-stage turbine. This design can make the turbine have good performance both on the ground and at high altitude. In addition, to ensure that the Mach numbers at the outlet of the first-stage guide vane 131 and the outlet of the first-stage rotor 121 of the first-stage turbine are equivalent and avoid strong shock wave losses caused by excessive local velocity; control the energy reaction degree of the first-stage turbine at about 0.5-0.6, which can make the Mach numbers at the outlet of the first-stage guide vane 131 and the outlet of the first-stage rotor 121 both near the speed of sound. To reduce the exhaust velocity of the turbine and reduce the exhaust loss, control the energy reaction degree of the second-stage turbine at about 0.45-0.50, which can make the gas do more work in the second-stage guide vane 132 and less work in the second-stage rotor 122, so as to reduce the Mach number at the outlet of the second-stage rotor 122. This design can effectively reduce the exhaust loss.

[0047] Specifically, Step 2 includes: when selecting the load coefficient, since the turbine efficiency is given, the turbine work L is given, and the turbine speed is determined by the overall and compressor, and the load coefficient is only related to the tangential velocity u. Therefore, when the load coefficient is given, the flow channel height of the turbine can be determined. When designing this turbine, increase the load coefficient of the first-stage turbine and decrease the load coefficient of the second-stage turbine, which can prevent the load of the second-stage turbine from being too large at high altitude and keep the turbine performance.

[0048] Among them, the load coefficient: μ = L / u 2The larger the load coefficient is, the greater the work done by the turbine at this stage accounts for the total work done by the entire turbine. Too large a load coefficient is detrimental to turbine performance and is generally controlled between 1.0-1.6.

[0049] Flow coefficient: The flow coefficient represents the flow level of airflow in the blade channel. The larger the flow coefficient, the more flow needs to pass through under the same geometric area, and the greater the airflow speed is, so it also needs to be controlled within an appropriate range.

[0050] Energy reaction: The energy reaction force represents the proportion of the work done by the rotor in this stage of the turbine to the work done by the entire stage of the turbine.

[0051] Where L is the turbine work, u is the tangential velocity of the rotor blade outlet, C1 a is the axial velocity of the guide outlet, C2 a Axial velocity of the moving blade outlet, p 1 is the static pressure at the guide vane outlet, p 2 is the static pressure at the rotor blade outlet, P 0 is the total pressure at the guide vane inlet, and k is the specific heat ratio.

[0052] For the two-stage transonic turbine, the design indexes are shown in Table 1. The parameters in Table 1 are only an embodiment of the present invention and are not intended to further limit the parameters.

[0053]

[0054]

[0055] Table 1. Design indicators

[0056] According to the design method introduced in the present invention, zero-dimensional parameter analysis and inter-stage parameter allocation are carried out on the turbine, as shown in Table 2.

[0057] Parameter First-stage turbine Second-stage turbine Flow rate kg / s 2.5 2.6 Expansion ratio 2.9 2.6 Adiabatic efficiency 0.86 0.865 Inlet total temperature K 1339 1057 Inlet total pressure MPa 0.806 0.28 Flow coefficient 0.56 0.70 Load coefficient 1.6 1.2 Energy reaction degree 0.56 0.46 Rotor outlet relative Mach number 1.07 1.03 Outlet absolute Mach number 0.55 0.65 Turbine outlet gas flow angle 62.3 87.8

[0058] Table 2. One-dimensional parameter allocation

[0059] As an embodiment of the present invention, computational fluid dynamics software is used to perform numerical simulation on the design scheme, and the turbine performance curve obtained is as follows: Figure 7As shown, it can be seen that when the turbine expansion ratio is 7.54, the adiabatic efficiency of the turbine is 0.879, the test efficiency reaches 0.885, and the turbine inlet flow rate also reaches 2.5 kg / s. All the design results meet the design requirements, verifying the feasibility of the scheme. In addition, the performance of the turbine under the condition of 13 km altitude is calculated. At the high-altitude design point, the turbine efficiency drops by about 2.7 percentage points compared with that on the ground, still meeting the high-altitude index of 0.845. While adopting the traditional design idea, the turbine performance at high altitude drops by 4.2 percentage points, not meeting the index requirements. Thus, it can be seen that this scheme not only meets the design index on the ground, but also has good performance at high altitude, having obvious advantages over the conventional turbine design method.

[0060] The design idea of the present invention that takes into account the high-altitude performance can be applied not only to the two-stage transonic turbine, but also to the two-stage and three-stage subsonic / transonic turbines, especially when the engine has the requirement of high-altitude use. The design idea of the present invention can improve the turbine performance of the engine at high altitude, enhance the power and power-to-weight ratio of the engine at high altitude, and broaden the application range of the engine, having broad application prospects.

[0061] Although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that: they can still modify the technical solutions recorded in the foregoing embodiments, or perform equivalent replacements for some of the technical features; and these modifications or replacements 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. An engine turbine with high-altitude performance, the engine turbine is arranged in an annular passage (110) at the end of a flame tube (100) of a combustion chamber (1) of the engine. The annular passage (110) is centered on the engine axis, with the side closer to the axis being set as an inner flow passage (112), and the side farther from the axis being set as an outer flow passage (111). Characterized in that the engine turbine includes a first-stage turbine, a second-stage turbine, an inner flow passage (112) and an outer flow passage (111). The first-stage turbine and the second-stage turbine are evenly distributed along the radial direction of the annular passage (110). Among them, the first-stage turbine includes a first-stage rotor (121) and a first-stage stator (131), the second-stage turbine includes a second-stage rotor (121) and a second-stage stator (132), and the first-stage stator (131), the first-stage rotor (121), the second-stage stator (132) and the second-stage rotor (121) are arranged in sequence from the side closer to the flame tube (100) to the side farther from the flame tube (100); the rotor blades of the first-stage rotor (121) and the second-stage rotor (121) include a blade root and a blade tip extending along the rotor blade body. The direction of the blade tip rotates clockwise relative to the direction of the blade root, and the bend-twist angle between the blade tip and the blade root is 25 - 35°; the intersection line between the section on the side closer to the flame tube (100) of the second-stage stator (132) and the body of the second-stage stator (132) is set as a second leading-edge profile (1321), and the included angle B between the second leading-edge profile (1321) and the outer flow passage (111) is 90°; the energy reaction degree of the first-stage turbine is 0.5 - 0.6, and the energy reaction degree of the second-stage turbine is 0.45 - 0.

5.

2. The engine turbine according to claim 1, Characterized in that during the process of the blade root of the first-stage rotor (121) and the second-stage rotor (121) extending towards the blade tip, the rotor blade body gradually narrows.

3. The engine turbine according to claim 1, Characterized in that during the process of the blade root of the first-stage stator (131) and the second-stage stator (132) extending towards the blade tip, the stator blade body gradually widens.

4. The engine turbine according to claim 1, Characterized in that a blade tip groove is provided on the blade tip end face of the rotor blade of the first-stage rotor (121).

5. The engine turbine according to claim 1, Characterized in that the intersection line between the section on the side closer to the flame tube (100) of the first-stage stator (131) and the body of the first-stage stator (131) is set as a first leading-edge profile (1311), and the included angle A between the first leading-edge profile (1311) and the vertical direction is 5 - 10°.

6. The engine turbine according to any one of claims 1 - 5, Characterized in that the annular passage (110) at the end of the second-stage rotor (121) is set as an expanding type, and the airflow passage area between the outer flow passage (111) and the inner flow passage (112) on the side farther from the flame tube along the airflow direction gradually increases.

7. A design method for an engine turbine according to any one of claims 1 - 6, It is characterized in that the design method includes: Zero-dimensional parameter analysis: Determine the load coefficient, flow coefficient, and energy reaction degree of the first-stage turbine and the second-stage turbine respectively according to the design indicators of the engine turbine; One-dimensional parameter calculation: Obtain the aerodynamic parameters and geometric parameters of the turbine blades of the first-stage turbine and the second-stage turbine of the engine respectively according to the basic principles of turbine aerodynamics; Three-dimensional modeling design: Calculate the inlet and outlet aerodynamic parameters and geometric parameters of the first-stage turbine and the second-stage turbine according to the load coefficient, flow coefficient, energy reaction degree, aerodynamic parameters, and geometric parameters, and perform three-dimensional modeling design of the engine turbine.

8. The design method according to claim 7, It is characterized in that the design indicators of the engine turbine include the flow rate index, expansion ratio index, and adiabatic efficiency index of the first-stage turbine and the flow rate index, expansion ratio index, and adiabatic efficiency index of the second-stage turbine.

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