Rotor structure capable of inhibiting mainstream high-temperature fuel gas from invading disc cavity and turbine

Abstract

The rotor structure comprises a rotor blade, the rotor blade comprises a front edge, a tail edge, a pressure molded surface and a suction molded surface, the front edge is an arc surface protruding outwards, the pressure molded surface and the suction molded surface are a concave surface and a convex surface on the rotor blade respectively, and the convex surface is a convex surface. A flange is arranged outside the blade surface at the joint of the front edge and the turbine moving blade hub, the outer edge of the flange is a smooth arc-shaped molded line, and the flange forms a three-dimensional belt-shaped front flange of the rotor blade; the flow characteristic of airflow at the front edge of the blade can be optimized through the design of the flange at the front edge of the blade, when mainstream high-temperature fuel gas flows through the front edge, the flange arranged outside the front edge can guide the airflow to smoothly bypass the blade, separation of the airflow at the front edge and generation of vortex are reduced, and therefore the possibility that the fuel gas invades a disc cavity is reduced; in addition, the aerodynamic performance and the overall reliability and stability of the turbine can be improved by applying the rotor structure.

Description

Technical Field

[0001] This application relates to the field of gas turbine technology, and more particularly to a gas turbine turbine, specifically to a rotor structure capable of suppressing the intrusion of mainstream high-temperature gas into the disk cavity and a turbine equipped with the rotor structure. Background Technology

[0002] In gas turbines, the turbine component is the key link in converting thermal energy into mechanical energy, and its performance directly affects the efficiency and thrust of the entire engine. With the continuous advancement of aero-engine technology, improving the thrust-to-weight ratio has become an important goal for enhancing engine performance, and increasing the turbine inlet gas temperature is the most direct and effective means to achieve this goal. Studies have shown that for every 100°C increase in turbine inlet gas temperature, engine thrust can increase by about 8%. However, the rate of increase in turbine inlet gas temperature is rapid, currently increasing at a rate of about 20-30°C per year, which far exceeds the high-temperature resistance limit of existing turbine materials. For example, the turbine inlet gas temperature of modern aero-engines has reached over 1500°C, while the melting point of turbine blade materials is typically only around 1100°C. This makes the turbine's operating environment extremely harsh, especially the high-speed rotating turbine disk and rotor blade root areas, which face the combined effects of multiple complex factors such as high temperature, high pressure, centrifugal load, vibration load, thermal stress, gas corrosion, and high-temperature oxidation.

[0003] To ensure the normal operation and sustained high-temperature resistance of turbine components in such harsh environments, effective cooling and thermal protection measures are crucial. Currently, this is mainly achieved by drawing cooling gas from the compressor and delivering it to the turbine disk cavity, cooling the turbine disk cavity and blades via the disk, blades, and sealing devices. During its entry into the disk cavity, the cooling gas not only absorbs and removes a significant amount of heat, reducing the temperature of the turbine components, but also forms a film on the blade surface, isolating the high-temperature combustion gases from direct contact with the blades, thus effectively protecting them. Simultaneously, the cooling gas entering the disk cavity also prevents the mainstream high-temperature combustion gases from intruding into the turbine disk cavity through the sealing structure. Current research generally suggests that the non-uniformity of the mainstream circumferential pressure caused by unsteady rotation and static conditions is the primary reason for the intrusion of high-temperature combustion gases into the disk cavity. Once these gases intrude, the turbine disk temperature becomes excessively high, severely impacting engine operating safety and service life. Furthermore, the sealing gas eventually enters the mainstream channel and interferes with the mainstream flow. Therefore, while increasing the sealing flow rate can improve sealing performance to some extent and prevent combustion gas intrusion, it will significantly reduce turbine aerodynamic efficiency. This is because the interference between the sealing gas and the mainstream gas will cause airflow disturbance and energy loss, increasing turbine aerodynamic losses.

[0004] Currently, improving rim sealing performance is an effective way to enhance sealing performance through sealing structure design or optimizing the mainstream flow field. Specific solutions for improving sealing performance by modifying the sealing structure design to better adapt to airflow pressure changes include the following:

[0005] 1. Honeycomb-Grate Seal Structure: This structure includes a rotating shaft, a honeycomb bushing, and sealing grates. The inner annular surface of the honeycomb bushing is composed of multiple honeycomb cavities, which are tilted at a certain angle towards the airflow direction, increasing the airflow's ability to penetrate the honeycomb cavity. Adjacent honeycomb cells are connected through internal jet holes. When airflow enters the honeycomb cavity, the higher-pressure airflow from the upstream honeycomb cavity enters the downstream honeycomb cavity through circular through-holes. The vortex in the downstream honeycomb cavity is compressed, generating a kidney-shaped vortex, which enhances mixing and dissipation. By optimizing the honeycomb structure, it strengthens the airflow mixing and energy dissipation inside the grate seal, further reducing fluid leakage.

[0006] 2. Improvement of the grate sealing structure: The combination of helical teeth and high steps increases the flow rate of the sealing cavity due to the increased inclination, which in turn increases the stagnation area of ​​the airflow at the knife-shaped teeth and increases the curvature of the streamline around the tooth tip; the airflow channel formed by the high step cross section is more tortuous, and the flow rate also increases, which increases the turbulent viscous loss. Compared with the reference structure, the leakage can be reduced by 17%.

[0007] 3. Turbine disk cavity sealing structure with pre-swirl induced airflow: Several pre-swirl induced airflow paths are set on the inner support ring, which are composed of induced airflow holes, air collection chambers and air outlets in sequence. After adopting the pre-swirl induced airflow path, the airflow velocity and pre-swirl angle of the sealed cold air outlet are increased, which can reduce the velocity triangle difference with the mainstream gas flow at the guide vane root outlet. This can significantly reduce the secondary flow loss in the moving blade passage. Compared with the existing disk cavity sealing structure, the turbine efficiency can be improved by nearly 1% after adopting this structure.

[0008] 4. Grate sealing device based on the principle of counter-current: An inwardly concave annular counter-current groove is added at each step to counter-current with the jet between the original teeth, reducing the air permeability of the grates and achieving the purpose of optimizing the sealing; at the same time, the space of the counter-current groove of appropriate size allows the dissipation vortex behind the teeth to develop fully, further aggravating the energy dissipation, which can significantly reduce the flow rate and improve the sealing effect.

[0009] The above-mentioned solutions improve the sealing effect, reduce gas leakage, and ensure the safe and efficient operation of turbine components by optimizing the geometry of the sealing structure, increasing airflow mixing and energy dissipation, and adjusting airflow speed and direction.

[0010] However, during turbine operation, there is relative motion between the rotor blades and stator blades. This relative motion causes periodic pressure fluctuations in the airflow on the blade surfaces. When these pressure fluctuations are transmitted to the main flow, they create a non-uniform circumferential pressure distribution within the main flow. This non-uniform pressure field makes it easier for the main flow combustion gas to penetrate from lower pressure areas into the turbine cavity. Therefore, reducing the non-uniformity of the main flow circumferential pressure to control combustion gas intrusion is a pressing technical problem that needs to be solved in this field, and it is necessary to propose a new structural technology solution to address this issue. Utility Model Content

[0011] This application provides a rotor structure and turbine that can suppress the intrusion of mainstream high-temperature gas into the disk cavity, thereby solving the problem that mainstream high-temperature gas in existing turbines easily intrudes into the disk cavity.

[0012] To achieve the above objectives, this application provides the following technical solution:

[0013] On one hand, this application provides a rotor structure capable of suppressing the intrusion of mainstream high-temperature gas into the disk cavity, including rotor blades. The rotor blades have a first end and a second end respectively formed opposite to each other along their height extension direction. The first end is a free end at the blade tip, and the second end is a fixed end at the blade root for connecting with the turbine impeller hub. The rotor blade includes a blade leading edge, a blade trailing edge, a blade pressure profile, and a blade suction profile. The blade leading edge and the blade trailing edge are respectively outwardly protruding arc surfaces. The blade pressure profile is a concave surface on the rotor blade, and the blade suction profile is a convex surface on the rotor blade. The blade pressure profile and the blade suction profile are arranged opposite to each other. A flange is provided on the outer surface of the blade at the connection between the blade leading edge and the turbine impeller hub. The outer edge of the flange is a smooth curved arc shape. The flange constitutes the three-dimensional strip-shaped leading flange of the rotor blade.

[0014] Furthermore, in the above technical solution, the rotor blade is a long strip-shaped sheet structure, the end face shape of the first end of the rotor blade is the same as the end face shape of the second end, and the end face size of the first end of the rotor blade is the same as the end face size of the second end.

[0015] Furthermore, the profiles of the blade pressure profile and the blade suction profile are each composed of multiple arc segments that are smoothly connected in sequence.

[0016] Furthermore, the end face shape of the fixed end of the blade root is formed by connecting the first curve, the second curve, the third curve, the fourth curve, and the fifth curve sequentially in a clockwise direction. The first curve and the second curve are connected to form an L-shaped curve, and the fourth curve and the fifth curve are connected to form an S-shaped curve. The rotor blade forms the blade suction profile corresponding to the blade surface of the L-shaped curve, the rotor blade forms the blade pressure profile corresponding to the blade surface of the S-shaped curve, and the rotor blade forms the blade trailing edge corresponding to the blade surface of the third curve. The L-shaped curve and the S-shaped curve are arranged opposite to each other. The bending direction of the L-shaped curve is the same as the bending direction of the fourth curve and opposite to the bending direction of the fifth curve.

[0017] Furthermore, the first curve and the fifth curve are arranged opposite each other, the second curve and the fourth curve are arranged opposite each other, and the third curve is used to connect the second curve and the third curve with an arc transition; the connection between the first curve and the second curve forms a smooth corner of the L-shaped curve, and the fourth curve and the fifth curve are smoothly connected.

[0018] Furthermore, the connection point of the first curve and the fifth curve forms the highest point of the leading edge of the blade. The horizontal straight line passing through the highest point is designated as the reference horizontal line. The straight line length of the first curve along the direction of the reference horizontal line is the straight line extension length of the three-dimensional strip-shaped leading flange on the suction surface of the blade, and this straight line extension length is designated as 'a'. The straight line length of the fifth curve along the direction of the reference horizontal line is the straight line extension length of the three-dimensional strip-shaped leading flange on the pressure surface of the blade, and this straight line extension length is designated as 'b'.

[0019] Furthermore, a represents 30% of the axial chord length of the rotor blade; b represents 30% of the axial chord length of the rotor blade.

[0020] Furthermore, the height of the flange along the direction from the leaf root to the leaf tip is the shaping height of the three-dimensional strip-shaped front flange. The shaping height of the front edge of the three-dimensional strip-shaped front flange is denoted as h, and the forward extension length of the flange relative to the front edge of the blade is denoted as c. c is 7.5% of the axial chord length of the rotor blade, and h is 7% of the blade height of the rotor blade.

[0021] Furthermore, the axial profile curve of the three-dimensional strip-shaped front flange in a two-dimensional rectangular coordinate system is expressed as follows:

[0022] x = R[1 - cos(2πt)]sin(2πt)

[0023] y = cos(2πt)

[0024] Where t∈[0,1], R represents the shape factor, and the value of R ranges from 0 to 1.

[0025] Furthermore, R = 0.5.

[0026] On the other hand, this application provides a turbine with the rotor structure described above that can suppress the intrusion of mainstream high-temperature gas into the disk cavity, including a moving impeller hub, a guide vane hub, rotor blades, guide vanes, and an axial disk cavity sealing structure; the blade roots of the rotor blades are connected and fixed to the moving impeller hub, the blade roots of the guide vanes are connected and fixed to the guide vane hub, the moving impeller hub and the guide vane hub are arranged opposite to each other, and the axial disk cavity sealing structure is provided at the connection between the moving impeller hub and the guide vane hub.

[0027] Compared with the prior art, this application has at least the following beneficial effects:

[0028] This application provides a rotor structure capable of suppressing the intrusion of mainstream high-temperature combustion gases into the turbine disk cavity. A flange is provided on the outer surface of the blade at the connection between the blade leading edge and the turbine impeller hub. The outer edge of the flange is a smooth curved line, forming a three-dimensional strip-shaped leading flange of the rotor blade. Through the above-mentioned three-dimensional strip-shaped leading flange design, the flow characteristics of the airflow at the blade leading edge are optimized, so that the airflow forms a more stable flow structure near the blade leading edge. When the mainstream high-temperature combustion gases flow through the blade leading edge, the flange provided on the outer edge can guide the airflow to bypass the blade more smoothly, reducing the separation of airflow and the generation of eddies at the leading edge, thereby reducing the possibility of combustion gases intruding into the disk cavity. In addition, the rotor blades in this application can more effectively prevent the intrusion of mainstream high-temperature combustion gases and improve the sealing efficiency of the turbine disk cavity. This not only helps to protect the components inside the turbine disk cavity from the corrosion of high-temperature combustion gases and extend their service life, but also reduces the thermal stress and harmful deformation caused by combustion gas intrusion, improving the overall reliability and stability of the turbine. Attached Figure Description

[0029] To more clearly illustrate the technical solutions in the embodiments of this application, the accompanying drawings used in the embodiments will be briefly described below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort. It should be understood that the specific shapes and structures shown in the drawings should not generally be regarded as limiting conditions for implementing this application; for example, based on the technical concepts disclosed in this application and the exemplary drawings, those skilled in the art are able to easily make conventional adjustments or further optimizations to the addition / reduction / classification, specific shapes, positional relationships, connection methods, and size ratios of certain units (components).

[0030] Figure 1This is a schematic diagram of the structure of a first-stage turbine with a three-dimensional strip-shaped front flange provided in this application, as one embodiment.

[0031] Figure 2 for Figure 1 A magnified view of a portion of point A in the middle;

[0032] Figure 3 This is a schematic diagram of the outline of a three-dimensional strip-shaped front flange of a rotor blade in one embodiment. The geometric design control parameters of the three-dimensional strip-shaped front flange are marked in the figure.

[0033] Figure 4 This is a schematic diagram of the longitudinal structure of a rotor blade in one embodiment, mainly illustrating the longitudinal structure from the blade root to the blade tip. The geometric design control parameters of the three-dimensional strip-shaped front flange are marked in the figure.

[0034] Figure 5 Axial cross-sectional curves of the three-dimensional strip-shaped front flange corresponding to different shape factors.

[0035] Explanation of reference numerals in the attached figures:

[0036] 1. Guide vane; 2. Rotor blade; 3. Three-dimensional strip-shaped front flange; 4. Guide vane hub; 5. Moving vane hub; 6. Axial disk cavity sealing structure; 7. First curve; 8. Second curve; 9. Third curve; 10. Fourth curve; 11. Fifth curve; 12. Smooth curved profile. Detailed Implementation

[0037] The present application will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0038] In the description of this application, expressions such as "comprising," "including," and "having" also imply "not limited to" (certain units, components, materials, steps, etc.). Terms such as "upper," "lower," "left," and "right" used in this application are generally for ease of intuitive understanding with reference to the accompanying drawings and are not absolute limitations on the positional relationships in the actual product. Changes in these relative positional relationships, without departing from the technical concept disclosed in this application, should also be considered within the scope of the description in this application.

[0039] Example 1

[0040] Improving the sealing performance of the turbine rim is an effective way to enhance sealing performance through sealing structure design or optimization of the mainstream flow field. Specific solutions for optimizing the mainstream flow field to make the airflow inside the turbine more uniform and stable, and to reduce the effects of unsteady interference, include the following:

[0041] 1. Non-axisymmetric endwall design optimization: By altering the geometry of the endwall, it is no longer a traditional axisymmetric structure but is optimized according to airflow characteristics. For example, grooves, protrusions, or other irregular shapes are added to the endwall to guide the flow direction and velocity distribution of the airflow. Non-axisymmetric endwalls can disrupt the original streamline structure, causing more vortices and mixing of the airflow near the endwall, thereby reducing the intensity and range of secondary flows. Simultaneously, this design can also change the flow path of the airflow within the blade passage, making it more evenly distributed on the blade surface, reducing the velocity and pressure gradients in local areas. The optimized non-axisymmetric endwall can significantly reduce aerodynamic losses inside the turbine and improve airflow efficiency and stability.

[0042] 2. Designing turbine blades using free deformation technology: Free deformation technology allows for parametric modeling and optimization of turbine blade geometry. By adjusting the position and number of control points, geometric parameters such as blade profile, camber, and twist can be flexibly altered. The optimized blade shape better adapts to airflow characteristics, resulting in smoother and more stable airflow on the blade surface, reducing airflow separation and vortex generation. Furthermore, adjusting aerodynamic parameters such as angle of attack and chord length can optimize the airflow velocity and pressure distribution within the blade passage. Turbine blades optimized using free deformation technology can improve aerodynamic efficiency while reducing the effects of unsteady airflow interference.

[0043] Based on the above-mentioned schemes for optimizing the mainstream flow field, this application provides an innovative rotor structure that can suppress the intrusion of mainstream high-temperature gas into the disk cavity. The rotor structure has a teardrop-shaped rotor leading edge. This application reduces the non-uniformity of the mainstream circumferential pressure by designing the teardrop-shaped rotor leading edge, thereby achieving the purpose of controlling gas intrusion.

[0044] Specifically, the rotor structure includes rotor blades 2, each blade having a first end and a second end arranged opposite to each other along its height extension direction. The first end is a free end at the blade tip, and the second end is a fixed end at the blade root for connection with the turbine impeller hub 5. Figure 1 The basic structure of the rotor blade 2 in this application is the same as that of the existing rotor blade, specifically including a blade leading edge, a blade trailing edge, a blade pressure profile, and a blade suction profile. The blade leading edge and the blade trailing edge are respectively outwardly protruding arc surfaces. The blade pressure profile is a concave surface on the rotor blade, and the blade suction profile is a convex surface on the rotor blade. The blade pressure profile and the blade suction profile are generally arranged opposite to each other.

[0045] In this application, a flange is provided on the outer surface of the blade at the connection between the leading edge of the blade and the turbine impeller hub. The outer edge of the flange is a smooth curved line, and the flange constitutes the three-dimensional strip-shaped leading flange 3 of the rotor blade 2.

[0046] The blade structure curve design in this application gives the fixed end of the blade root specific geometric features, which can effectively disrupt and change the flow path and velocity distribution of the airflow near the blade root, increasing the flow resistance and turbulence. When the mainstream high-temperature combustion gas attempts to invade the disk cavity through the gap between the blade root and the turbine impeller hub 5, it will be hindered by this complex curve shape, and the flow direction and velocity of the airflow will be changed multiple times, thereby reducing the intrusion capability of the combustion gas. In addition, the design of the three-dimensional ribbon-shaped leading flange 3 of the rotor blade 2 can optimize the flow characteristics of the airflow at the blade leading edge, so that the airflow forms a more stable flow structure near the blade leading edge. When the mainstream high-temperature combustion gas flows through the blade leading edge, the three-dimensional ribbon-shaped leading flange 3 can guide the airflow to bypass the blade more smoothly, reducing the separation of the airflow and the generation of vortices at the leading edge, thereby reducing the possibility of combustion gas intruding into the disk cavity.

[0047] In a preferred embodiment of this application, the rotor blade 2 is a long strip-shaped sheet structure. The end face shape of the first end of the rotor blade 2 is the same as the end face shape of the second end, and the end face dimensions of the first end of the rotor blade 2 are the same as the end face dimensions of the second end. Furthermore, the profiles of the blade pressure profile and the blade suction profile are respectively formed by multiple arc segments that are smoothly connected in sequence, such as... Figure 1 .

[0048] In one specific embodiment of this application, the end face shape of the fixed end of the rotor blade 2 is formed by connecting the first curve 7, the second curve 8, the third curve 9, the fourth curve 10, and the fifth curve 11 sequentially end to end in a clockwise direction, as shown below. Figure 3 Wherein: the first curve 7 and the second curve 8 form an L-shaped curve; the fourth curve 10 and the fifth curve 11 form an S-shaped curve; the rotor blade 2, corresponding to the blade surface of the L-shaped curve, forms the blade suction profile; the rotor blade 2, corresponding to the blade surface of the S-shaped curve, forms the blade pressure profile; and the rotor blade 2, corresponding to the blade surface of the third curve, forms the blade trailing edge. See also Figure 3 The L-shaped curve and the S-shaped curve are set opposite to each other. The bending direction of the L-shaped curve is the same as that of the fourth curve 10 and opposite to that of the fifth curve 11. In this application, the three-dimensional strip-shaped front flange 3 is designed by superimposing a smooth curved line 12 on the leading edge of the blade. That is, a strip-shaped flange is provided outside the leading edge of the blade, and the outer edge of the flange is a smooth curved line 12. This flange constitutes the three-dimensional strip-shaped front flange 3 of the rotor blade 2. The design process and design parameters of the three-dimensional strip-shaped front flange 3 are as follows: Figure 3 , 4 .

[0049] See also Figure 3The first curve 7 and the fifth curve 11 are set opposite to each other, forming the leading edge of the blade. The second curve 8 and the fourth curve 10 are set opposite to each other. The third curve 9 is used to connect the second curve 8 and the third curve 9 with an arc transition. The connection between the first curve 7 and the second curve 8 forms a smooth corner of an L-shaped curve. The fourth curve 10 and the fifth curve 11 are smoothly connected.

[0050] In this application, the design parameters of the three-dimensional strip-shaped anterior flange 3 are as follows: Figure 3 , 4 Specifically, as indicated in the diagram, the connection point of the first curve 7 and the fifth curve 11 forms the highest point of the blade's leading edge. The horizontal line passing through this highest point is designated as the reference horizontal line. The length of the first curve 7 along the reference horizontal line is the linear extension length of the three-dimensional strip-shaped leading flange on the blade's suction surface, denoted as 'a'. The length of the fifth curve 11 along the reference horizontal line is also the linear extension length of the three-dimensional strip-shaped leading flange on the suction surface, denoted as 'b'. The height of the flange from the blade root to the blade tip is the design height of the three-dimensional strip-shaped leading flange, denoted as 'h'. The forward extension length of the flange relative to the blade's leading edge is denoted as 'c'. Figure 4 The middle arrow Q indicates the direction of the mainstream high-temperature gas flow.

[0051] In one specific design example, a and b in this application are the same, both being 30% of the axial chord length of the rotor blade; c is 7.5% of the axial chord length of the rotor blade, and h is 7% of the rotor blade height. Of course, in other design examples, the values ​​of a and b can be customized according to application requirements.

[0052] In a preferred embodiment of this application, after determining the design parameters of the three-dimensional strip-shaped leading flange of the blade in the upper end region, the smooth curved profile 12 is combined with the leading edge of the blade to obtain the three-dimensional strip-shaped leading flange. The axial section curve of the three-dimensional strip-shaped leading flange in this application is expressed as a function in a two-dimensional rectangular coordinate system as follows:

[0053] x = R[1 - cos(2πt)]sin(2πt)

[0054] y = cos(2πt)

[0055] Where t∈[0,1], R represents the shape factor, and the value of R ranges from 0 to 1.

[0056] In a specific design example, R can take values ​​of 0, 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1, etc. Figure 5The teardrop-shaped profiles are shown for different shape factors. Calculations show that R = 0.5 is preferred in this application. Numerical calculations demonstrate that the turbine device with a teardrop-shaped rotor leading edge designed in this application reduces the non-uniformity of the mainstream circumferential pressure compared to the original turbine device, resulting in improved sealing performance. At low sealing flow rates, the sealing efficiency is improved by 7.35%.

[0057] Example 2

[0058] Based on the rotor structure provided in Embodiment 1 above, which can suppress the intrusion of mainstream high-temperature gas into the disk cavity, this embodiment provides a turbine equipped with this rotor structure. See [link to previous embodiment]. Figure 1 The turbine includes a moving impeller hub 5, a guide vane hub 4, rotor blades 2, guide vanes 1, and an axial disk sealing structure 6. The blade root of the rotor blade 2 is connected and fixed to the moving impeller hub 5, and the blade root of the guide vane 1 is connected and fixed to the guide vane hub 4. The moving impeller hub 5 and the guide vane hub 4 are arranged opposite to each other, and an axial disk sealing structure 6 is provided at the connection between the moving impeller hub 5 and the guide vane hub 4.

[0059] This application, through the aforementioned unique blade structure curve design and three-dimensional strip-shaped front flange design, enables the rotor blades to more effectively prevent the intrusion of mainstream high-temperature combustion gases, thereby improving the sealing efficiency of the turbine disk cavity. This not only helps protect the components within the turbine disk cavity from the erosion of high-temperature combustion gases, extending their service life, but also reduces thermal stress and harmful deformation caused by combustion gas intrusion, improving the overall reliability and stability of the turbine. Furthermore, while improving the sealing effect, this design also optimizes the turbine's aerodynamic performance. By rationally guiding and controlling the airflow, it reduces airflow losses and disturbances, improves the turbine's energy conversion efficiency, and enables the turbine to operate more efficiently in high-temperature, high-pressure environments.

[0060] The technical features of the above embodiments can be combined in any way (as long as there is no contradiction in the combination of these technical features). For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described; these embodiments not explicitly written should also be considered to be within the scope of this specification.

[0061] The present application has been described in a relatively specific and detailed manner above through general descriptions and specific embodiments. It should be understood that, based on the technical concept of the present application, several conventional adjustments or further innovations can be made to these specific embodiments; however, as long as they do not depart from the technical concept of the present application, the technical solutions obtained by these conventional adjustments or further innovations also fall within the protection scope of the claims of the present application.

Claims

1. A rotor structure capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, characterized in that, The rotor blade includes rotor blades, each having a first end and a second end arranged opposite to each other along its height extension direction. The first end is a free end at the blade tip, and the second end is a fixed end at the blade root for connection with the turbine impeller hub. Each rotor blade includes a blade leading edge, a blade trailing edge, a blade pressure profile, and a blade suction profile. The blade leading edge and the blade trailing edge are respectively outwardly protruding arc surfaces. The blade pressure profile is a concave surface on the rotor blade, and the blade suction profile is a convex surface on the rotor blade. The blade pressure profile and the blade suction profile are arranged opposite to each other. A flange is provided on the outer surface of the blade at the connection between the blade leading edge and the turbine impeller hub. The outer edge of the flange is a smooth curved line, and the flange constitutes the three-dimensional strip-shaped leading flange of the rotor blade.

2. The rotor structure according to claim 1, capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, is characterized in that, The rotor blade has a long strip-shaped structure. The end face shape of the first end of the rotor blade is the same as the end face shape of the second end, and the end face size of the first end of the rotor blade is the same as the end face size of the second end. The profiles of the blade pressure profile and the blade suction profile are each composed of multiple arcs that are smoothly connected in sequence.

3. The rotor structure according to claim 1, capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, is characterized in that, The end face shape of the fixed end of the blade root is formed by connecting the first curve, the second curve, the third curve, the fourth curve, and the fifth curve in a clockwise direction. The first curve and the second curve are connected to form an L-shaped curve, and the fourth curve and the fifth curve are connected to form an S-shaped curve. The rotor blade forms the suction profile corresponding to the surface of the L-shaped curve, the pressure profile corresponding to the surface of the S-shaped curve, and the trailing edge corresponding to the surface of the third curve. The L-shaped curve and the S-shaped curve are arranged opposite to each other. The bending direction of the L-shaped curve is the same as the bending direction of the fourth curve and opposite to the bending direction of the fifth curve. The first curve and the fifth curve are arranged opposite each other, the second curve and the fourth curve are arranged opposite each other, and the third curve is used to connect the second curve and the third curve with an arc transition; The connection between the first curve and the second curve forms a smooth turn of the L-shaped curve, and the fourth curve and the fifth curve are smoothly connected.

4. The rotor structure according to claim 3, capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, characterized in that, The connection point of the first curve and the fifth curve forms the highest point of the leading edge of the blade. The horizontal straight line passing through the highest point is recorded as the reference horizontal line. The straight line length of the first curve along the direction of the reference horizontal line is the straight line extension length of the three-dimensional strip-shaped front flange on the suction surface of the blade. The straight line extension length of the three-dimensional strip-shaped front flange on the suction surface of the blade is recorded as a. The straight line length of the fifth curve along the reference horizontal line is the straight line extension length of the three-dimensional strip-shaped front flange on the blade pressure profile, and the straight line extension length of the three-dimensional strip-shaped front flange on the blade pressure profile is denoted as b.

5. The rotor structure according to claim 4, capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, is characterized in that, a is 30% of the axial chord length of the rotor blade; b is 30% of the axial chord length of the rotor blade.

6. The rotor structure according to claim 4, capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, is characterized in that, The height of the flange along the direction from the leaf root to the leaf tip is the shaping height of the three-dimensional strip-shaped front flange. The shaping height of the three-dimensional strip-shaped front flange is denoted as h, and the forward extension length of the flange relative to the leading edge of the blade is denoted as c. c is 7.5% of the axial chord length of the rotor blade, and h is 7% of the blade height of the rotor blade.

7. The rotor structure capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity according to claim 1 or 6, characterized in that, The axial profile curve of the three-dimensional strip-shaped front flange in a two-dimensional rectangular coordinate system is expressed as follows: x = R[1 - cos(2πt)]sin(2πt) y = cos(2πt) Where t∈[0,1], R represents the shape factor, and the value of R ranges from 0 to 1.

8. The rotor structure according to claim 7, capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, characterized in that, R＝0.5。 9. A turbine equipped with the rotor structure of claim 1, capable of suppressing the intrusion of mainstream high-temperature combustion gas into the disk cavity, characterized in that, It includes a moving impeller hub, a guide impeller hub, rotor blades, guide blades, and an axial disk cavity sealing structure; the blade roots of the rotor blades are connected and fixed to the moving impeller hub, the blade roots of the guide blades are connected and fixed to the guide impeller hub, the moving impeller hub and the guide impeller hub are arranged opposite to each other, and the axial disk cavity sealing structure is provided at the connection between the moving impeller hub and the guide impeller hub.

Images