Anti-rotation structure, use and verification method of wall plate applied to turbine blade inlet

Abstract

The present invention relates to a wall plate anti-swirl structure applied to the inlet of a turbine blade, comprising left and right blade surfaces, a non-reactive combustion chamber swirler, the inner wall of the non-reactive combustion chamber cavity, left and right periodic surfaces of the non-reactive combustion chamber cavity, and a non-reactive combustion chamber inlet section anti-swirl thin plate. By designing a thin plate structure at the inlet of the turbine blade, the residual swirl at the outlet of the combustion chamber can be effectively suppressed, and the outlet hot spot can be dispersed, the swirl intensity of the turbine blade inlet flow field can be significantly reduced, the uniformity of the turbine inlet flow field can be increased, the cooling efficiency of the turbine first-stage guide vane can be improved, the maximum temperature and temperature gradient of the first-stage guide vane can be reduced, and the amount of cooling gas and the difficulty of its cooling design can be reduced. The present invention sets an anti-swirl thin plate structure at the inlet of the first-stage guide vane, which has a simple structure and can be applied to the boundary area of ​​the turbine of the combustion chamber of an aircraft engine to control the swirl intensity.

Description

Technical Field

[0001] The present invention belongs to the technical field of gas turbine blade cooling, and in particular relates to a wall plate rotation-inhibiting structure applied to a turbine blade inlet and a method for verifying its use. Background Art

[0002] Aircraft engines are the crown of today's industry, representing a country's cutting-edge industrial design and manufacturing capabilities. As one of the most critical parameters of aircraft engines, turbine inlet temperature significantly impacts their overall performance. Increasing the inlet temperature not only boosts engine thrust but also improves cycle efficiency by adjusting the boost ratio. Raising the inlet temperature has become a constant design goal for engineers. Currently, the inlet temperature of advanced aircraft engines exceeds 2000K. This extreme heat exceeds the temperature resistance limit of the turbine material, necessitating the use of appropriate cooling technology. The current primary cooling solution combines internal convection and impingement cooling with external film cooling to provide thermal protection for turbine blades.

[0003] However, for the cooling structure design of the first-stage guide vane, due to the uneven distribution of the combustion chamber outlet temperature, a hot spot will be formed at the inlet. The local high-temperature area of ​​the hot spot makes the guide vane bear a higher temperature load, which will increase the risk of local high-temperature burning of the blade. At the same time, separation of cold and hot gases will occur in the moving blades, which brings great difficulties and uncertainties to the cooling design of the aircraft engine; at the same time, as the axial length of the engine combustion chamber gradually shortens, the swirl generated by the combustion chamber swirler has not completely dissipated at the outlet, and the residual swirl will enter the turbine component with the mainstream. The swirl will form an obvious secondary flow at the blade mid-span. In the process of propagating downstream, the strong swirl will suck in the nearby film cooling gas, and the secondary flow vortex will carry the cooling gas downward, resulting in insufficient cold air coverage on the blade surface, greatly reducing the cooling efficiency. Current research shows that the combined effect of swirl and hot spots will greatly affect the cooling efficiency of turbine blades. The hot spots themselves have higher temperatures, and the swirl weakens the thermal protection of the turbine blades by sucking in the cooling air film. At the same time, the interaction between the two is manifested in the following: the hot spots continue to propagate downstream under the entrainment of the swirl, and the high turbulence of the swirl enhances the heat transfer between the high-temperature fluid and the wall panel. The interaction between the two further worsens the thermal environment in which the turbine blades operate.

[0004] The current mainstream research focuses on exploring the flow field characteristics of the combustion chamber and turbine under the coupled relationship, so as to obtain the migration characteristics of swirl and hot spots in the flow field, and finally optimize the cooling design of the turbine blades accordingly. The focus is on: by designing the cooling structure or changing the upstream structure, serving the cooling design of the downstream guide vanes. At present, there are few structural modifications upstream of the guide vanes in academia and engineering. The only research was carried out by the Oxford University team. They proposed a combustion chamber-turbine integrated design, with a span-wise wall panel structure designed upstream of the guide vanes to reduce the heat load on the leading edge of the guide vanes. However, this did not weaken the swirl intensity of the flow field. The strong swirl was even more completely preserved, and the hot spots were transported downstream along the channel. However, the field of swirl control and hot spot control based on swirl and aerodynamic considerations is still very blank. There is no public structural design and research for the rectification of the guide vane inlet. Swirl control and hot spot control can make the turbine inlet flow field tend to be uniform, optimize the aerodynamic angle of attack, and ensure the reliability of the one-dimensional parameter design during the design process; more importantly, it can reduce the consumption of cooling gas and use less cooling gas to achieve the same cooling effect. The flow field control at the turbine inlet is of great significance to the aerodynamic and cooling design of the turbine blades.

[0005] In summary, how to design a swirl suppression structure that can suppress swirl and hot spots in the flow field at the combustion chamber outlet, significantly reduce the swirl intensity at the turbine inlet, improve the uniformity of the flow field at the turbine blade inlet, and ensure a simple and convenient structure, is a problem that needs to be solved urgently in this field. Summary of the Invention

[0006] To address the above problems, the present invention provides a wall plate vortex suppression structure for use at a turbine blade inlet, comprising a left blade surface, a right blade surface, a non-reactive combustion chamber swirler, a non-reactive combustion chamber cavity inner wall plate, a non-reactive combustion chamber cavity inner wall plate, left and right periodic surfaces of the non-reactive combustion chamber cavity, and a non-reactive combustion chamber inlet turbine inlet section vortex suppression thin plate. The upper and lower wall surfaces, left and right blade surfaces, inlet and outlet surfaces, and corresponding left and right periodic surfaces of the non-reactive combustion chamber enclose a combustion chamber and turbine fluid domain. The non-reactive combustion chamber swirler is located at the front side of the cavity. The swirl characteristics of the swirler are used to construct a swirl field. Cooling gas is then introduced through the non-reactive combustion chamber cavity inner and outer wall plates to construct a hot spot flow field. At the turbine inlet within the non-reactive combustion chamber cavity, the turbine inlet section vortex suppression plate is connected to the left and right blade surfaces to form a baffle area. To simplify modeling, the left and right periodic surfaces represent a model unit. The actual baffle area should be a circular ring structure, serving as the wall plate vortex suppression structure for use at the turbine blade inlet.

[0007] As a further improvement of the above technical solution: a plurality of periodic rectangular protrusions are provided on the surface of the vortex suppression thin-walled plate, corresponding to the thin plate and covering the surface thereof.

[0008] As a further improvement of the above technical solution, the radial position of the vortex suppression thin plate is reduced, and an identical vortex suppression thin plate is arranged on its upper side to form a double-plate vortex suppression structure.

[0009] As a further improvement of the above technical solution, a plurality of periodic wave ribs are provided on the surface of the vortex suppression wave plate to ensure the vortex suppression capability of the structure while reducing the flow loss of the structure.

[0010] The present invention also discloses a method for using a wall plate swirl suppression structure applied to a turbine blade inlet, comprising a wall plate swirl suppression structure applied to a turbine blade inlet, characterized in that when high-pressure gas at a compressor flows into a combustion chamber from a fluid domain inlet surface (1), the gas generates a swirl of a certain intensity after passing through a swirler surface (2) and a swirl cylinder channel surface (6), and the swirl suppression wall plate suppresses and controls the swirl, thereby reducing its swirl intensity and enhancing the unevenness of the turbine blade inlet, thereby improving the cooling design effect.

[0011] The present invention also provides a simulation and verification method for a wall plate swirl suppression structure applied at the inlet of a turbine blade. First, in order to simulate the swirl characteristics in actual engine operation, a non-reactive combustion chamber is designed upstream thereof. By reproducing aerodynamics and heat transfer, the flow field distribution under actual working conditions is accurately simulated. It adopts any of the above-mentioned wall plate swirl suppression structures applied at the inlet of a turbine guide vane. After being combined with the non-reactive combustion chamber, the high-temperature gas from the upstream combustion chamber flows out from the swirler outlet to form a swirl. After flowing through the swirl suppression wall plate, the swirl is restricted by the wall plate, the swirl structure is destroyed, the swirl of the flow field is weakened, and the uniformity of the flow field at the guide vane inlet is improved.

[0012] The advantages of the present invention are: by arranging a swirl suppression structure upstream of the turbine blade, the swirl in the upstream flow field is guided and suppressed, the non-flow directional rotational flow is weakened, and it is beneficial to the smooth uniformity of the guide vane inlet. The swirl suppression structure of the present invention has the characteristics of simple structure, easy processing and good cooling effect, and can be applied to the inlets of various turbine guide vanes. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] Figure 1 This is an axonometric drawing of the present invention.

[0014] Figure 2 It is a top view of the present invention.

[0015] Figure 3 It is a right side view of the present invention.

[0016] Figure 4 For the present invention Figure 2AA cross-sectional view.

[0017] Figure 5 It is the circumferential average swirl radial angle curve distributed along the blade height in a certain section of the flow field with a wall panel vortex suppression structure.

[0018] Figure 6 It is the circumferential average swirl tangential angle curve distributed along the blade height at a certain section in the flow field with a wall panel vortex suppression structure. DETAILED DESCRIPTION

[0019] In order to enable those skilled in the art to better understand the technical solution of the present invention, the present invention is described in detail below with reference to the accompanying drawings. The description in this part is merely exemplary and explanatory and should not have any limiting effect on the scope of protection of the present invention.

[0020] In order to make the purposes, features and advantages of the present invention more understandable, a detailed description of the specific embodiments of the present invention will be given below. In the following description, many specific details are described to help fully understand the present invention. However, the present invention can be implemented in many other ways different from the description, and those skilled in the art can make similar changes without violating the connotation of the present invention. Therefore, the present invention is not limited to the specific embodiments disclosed below.

[0021] Combined with attachment Figure 1-4 This paper provides a wall plate anti-swirl structure applied to the inlet of turbine blades. Since the anti-swirl structure is connected to the guide vane, its main function is to suppress the swirl at the outlet of the combustion chamber. In order to demonstrate the anti-swirl structure and reflect its main function, the front-end combustion chamber swirler is used as a fluid domain for illustration during the demonstration, and periodic boundary conditions are used to reduce the demonstration volume. The entire model includes the fluid domain inlet surface 1, the swirler surface 2, the fluid domain lower wall 3, the fluid domain right blade 4, the fluid domain upper wall 5, the swirl channel surface 6, the fluid domain periodic surface 7, the anti-swirl wall plate 8, and the fluid domain left blade 9; the anti-swirl wall plate 8 is connected to the fluid domain right blade 4 and the fluid domain left blade 9. The high-temperature fluid flows in from the fluid domain inlet surface 1, flows in the flow channel area formed by the fluid domain lower wall 3 and the fluid domain upper wall 5, generates a swirl after flowing through the swirler surface 2 and the swirl channel surface 6, and continues to be transmitted to the blade downstream.

[0022] The radius difference between the lower wall 3 and the upper wall 5 of the fluid domain in this model is 60 mm, which is consistent with the design height of the combustion chamber of the actual engine design. The structure is to reduce the display volume. A model unit is selected as 1 / 23 unit on a circumference. Figure 4The angle between the boundaries of the fluid unit on both sides is α, where α is 15.65°, indicating that 23 of the model units can form a complete annular fluid domain with a circumference of 360° and a completed annular anti-rotation wall shell through rotation array operation, that is, after periodic rotation supplementation, the anti-rotation wall plate 8 actually forms an annular anti-rotation wall shell.

[0023] The swirler surface 2 is used to construct the rotating flow field required in the combustion chamber. After flowing through the swirler surface 2, the fluid generates a swirl of a certain intensity. The swirl intensity is protected by the swirl channel surface 6 to prevent it from dissipating too quickly. The anti-swirl wall plate 8 is arranged in front of the blades 4 on the right side of the fluid domain and the blades 9 on the left side of the fluid domain, and is positioned approximately at the middle radius position between the lower wall surface 3 of the fluid domain and the upper wall surface 5 of the fluid domain.

[0024] In order to verify the anti-rotation structure mentioned in this article, a simulation verification method for the anti-rotation structure of the wall plate at the inlet of the turbine blade was also invented. The specific implementation is as follows:

[0025] 1. Calculate the aerodynamic parameters of the mixed flow field inlet according to the required working conditions. For example, the average temperature used in the simulation in this article is 450K, and the calculation is based on the enthalpy conservation of the fluid: m c ·T c +m h ·T h =m in ·T in (T represents the total temperature of the fluid, m represents the mass flow rate, subscript c represents the cooling gas, subscript h represents the high-temperature gas, and in represents the blade inlet.) The high-temperature and low-temperature gas flow distribution is fixed at 65:35, and the temperature ratio is 1.77. The temperature is determined. After the temperature is determined, the cyclone design is completed according to the working conditions. The formula S is used. N =2 / 3(1-(D SW,i / D SW,o ) 3 / 1-(D SW,i / D SW,o ) 2 ), estimate the design cyclone flow rate. N Indicates the degree of swirl, the letter D indicates the diameter, where SW is the abbreviation of cyclone, the subscript i indicates the inside, and the subscript o indicates the outside, for example D SW,i Indicates the inner diameter of the cyclone.

[0026] 2. Taking into account the overall considerations of processing accuracy and flow field loss control, the thickness of the anti-swirl baffle is controlled to a constant value. In essence, the simulation of the combustion chamber simulator requires the use of high-precision simulations such as LES (Largeeddy simulation) and SAS (Scaleadaptive simulation) to fully reproduce the flow field. However, the simulation requires a lot of computing resources. Although RANS (Reynolds average Navier Stockes simulation) cannot fully reproduce the flow field characteristics, it can capture the center of the swirl for verification and fitting empirical relationships. The radial position of the swirl baffle is determined by the radius of the swirl center and the radius of the inner wall plate. The radius of the swirl center is defined as r i , define the radius of the swirl center as r o , according to the empirical formula, select r = 1.2373r i +0.7831r o As the placement radius of the anti-rotation wall plate, the above empirical formula is derived as follows:

[0027] 1) A model with a combustion chamber simulator is selected for modeling and simulation, in which the anti-rotation wall panel radius is set at eleven radius positions, evenly distributed between the inner wall panel radius and the outer wall panel radius.

[0028] 2) Use RANS to simulate the wall panel anti-rotation structure at different radial positions.

[0029] 3) After the simulation is completed, the total pressure loss coefficient at the guide vane outlet is As an evaluation index of aerodynamic loss (where superscript * represents total pressure, subscript 1 represents inlet, subscript 2 represents outlet, and letter p represents pressure), the radial rotation angle α = arctan u r / u z and tangential rotation angle β = arctan u θ / u z (where u represents velocity, subscript z represents axial direction, r represents radial direction, and θ represents tangential direction) is used as an indicator of the swirl characteristics of the flow field.

[0030] 4) In this way, for the wall plate anti-rotation structure at the i-th radial position, it corresponds to three parameters α i , β i 、C pti , taking each group as a set of parameter pairs, where α i , β i It is a negative indicator (that is, the higher the parameter is, the more unfavorable it is for the final evaluation). After making it positive, all parameters are normalized and the average value of α is obtained. and variance σ(α), then use the parameter Unify the two. Similarly for β and C pt Use the same steps to obtain s(β) and s(C pti) Then, the three coefficients are normalized, that is, s(α)′=s(α) / s total s(β)′=s(β) / s total 、s(C pti) ′=s(C pti ) / s total , get the weight coefficient of each evaluation index, and the final evaluation formula m i =α i s(α)′ β i s(β)′ ·C pti s(Cpti)′ calculate.

[0031] 5) According to the aforementioned evaluation index m i Among the ten radial positions mentioned above, the fifth radial position m5 has the highest evaluation index, indicating that under the comprehensive aerodynamic loss and flow field rotation evaluation indicators, this radial position can achieve the purpose of anti-rotation to the greatest extent while ensuring a low level of aerodynamic loss increment.

[0032] 6) To verify the reliability of the empirical position, the original cyclone configuration was retained, but the radius of the circular axis center was adjusted to achieve the effect of changing the radius of the cyclone center. The aforementioned evaluation criteria were still used for evaluation. After adjusting the radius position four times, the optimal radius position showed a linear relationship with the height of the cyclone center position. Through curve fitting, the empirical relationship formula for the optimal radius position was obtained: r = 1.2373r i +0.7831r o , this empirical formula is used to select and determine the radial position of the anti-rotation wall plate.

[0033] 3. Install the swirler on the front side of the unreactive combustion chamber model. Cooling and mixing holes are opened in the inner and outer cavities of the unreactive combustion chamber. If the cooling and mixing holes are fully modeled, the huge amount of grids will put higher requirements on computational efficiency and storage space. In order to reduce the simulation time and the memory occupied by the simulation results, the air film hole structure of the wall panel where the divergent cooling hole is located is removed, and the entire wall panel is used as the inlet of the cooling air. In this process, the simplified principle adopted is the identity of the axial wall panel momentum. The parameters for the wall panel are defined: porosity σ = A holes / A plate =∑ holes (πD 2 / 4cos(α)) / A plate , where the jet angle α is given by the formula α=tan -1 (Un / U t ) is defined as the angle between the jet hole and the wall. The purpose of this definition is to characterize the momentum relationship between the simplified model and the original jet. The quantitative relationship between the two can be written as: and The original jet angle α is changed to α by the formula derivation mod , the two satisfy the equation: tan(α mod )=σtan(α), this theory transforms the original complex jet structure into a simple wall plate inlet and adjusts α to α mod In order to ensure the conservation of momentum near the jet, the uniform boundary assumption under porous divergent cooling is used to reduce the amount of data in the simulation calculation and facilitate subsequent verification.

[0034] 4. After combining the designed swirl suppression structure with the blades and then assembling it with the non-reactive combustion chamber, the high-temperature gas from the upstream combustion chamber forms a swirl after flowing out of the swirler outlet. The swirl suppression panel is actually located slightly below the center of the swirl. After flowing through the swirl suppression panel, the radial momentum of the swirl above the panel is restricted by the panel and weakened. The original large swirl structure is reduced to a small swirl structure above the panel. After the swirl below the panel is destroyed, a complete vortex structure cannot be formed on the flow surface. Overall, the swirl of the flow field is weakened. The swirl suppression panel's inhibitory effect is concentrated near the radius of the panel, and the flow field near the upper wall slightly away from the swirl suppression structure is less affected.

[0035] In order to specifically illustrate the size of the structure, a specific embodiment is used for illustration. In this embodiment, the structural parameters of the anti-rotation wall plate are mainly described by the wall plate thickness k and the wall plate width l. In this embodiment, the anti-rotation wall plate thickness k is 0.2 mm, which can be appropriately widened to 1 mm during actual application. The anti-rotation wall plate width l is 25 mm, which can be appropriately adjusted within 15 mm to 35 mm according to the combustion chamber length and the combustion chamber turbine interface length. The annular anti-rotation wall plate is connected to all the turbine first-stage guide vanes. The anti-rotation wall plate surface is smooth and has no other structures.

[0036] In this embodiment, high-temperature gas from the compressor flows from the fluid domain inlet surface 1, then through the swirler surface 2 and swirl channel surface 6, forming a vortex within the flow field and propagating downstream. The vortex significantly affects the heat exchange performance and temperature distribution on the surfaces of blades 4 on the right side of the downstream fluid domain and blades 9 on the left side of the fluid domain. The vortex entrains cooling gas, reducing the coverage of film cooling and indirectly generating hot spots that impact localized blades, hindering subsequent cooling design.

[0037] The design of the anti-swirl wall panel is to control the swirl of the mainstream flow field, reduce the aerodynamic impact of the swirl on the first-stage guide vanes, and weaken the migration of hot spots. The purpose of the design is to weaken, rather than control the swirl field but retain it, so it is necessary to destroy the swirl structure of the flow field, which makes the arrangement of the swirl wall panel adopt a circumferential setting. The radially arranged wall panel structure will also have an impact on the swirl, but in order to meet the needs of the blade flow channel, the wall panel will be set at the leading edge of the blade, which will eventually make the swirl pass completely through the middle flow channel of the wall panel, and the swirl intensity cannot be suppressed. The effect of the circumferential arrangement is to make the main flow field structure of the vortex destroyed due to the structure of the wall panel when the vortex passes through the flow channel, thereby reducing the vortex intensity. After determining that the structure is circumferentially arranged, the blades and the wall panel are designed to be integrated. One reason is that in the future engine manufacturing process, the blades will be processed by full-ring processing. The other reason is that the shedding vortex at the tail of the wall panel will have a worsening effect on the heat transfer characteristics of the blade leading edge. After the wall panel is moved backward, the shedding vortex falls directly into the flow field and will not affect the heat transfer characteristics of the blade leading edge. The existence of the anti-vortex wall panel 8, after the original vortex flows through this place, its vortex structure is destroyed by the wall panel, and the vortex cannot be fully developed. It can effectively reduce the vortex intensity, improve the flow field uniformity at the inlet of the turbine blade, and improve the cooling effect downstream.

[0038] Figure 5 The circumferential average radial swirl angle along the blade height at a certain section with and without the anti-swirl wall structure was compared. Figure 6 The circumferential average tangential swirl angle along the blade height at a specific cross section was compared for blades with and without the anti-swirl wall structure. It should be noted that the above simulations are based on a simulation verification method for a wall anti-swirl structure applied to the turbine blade inlet, and the verification process is feasible.

[0039] In this embodiment, to ensure comparability of the results, the flow conditions with and without the wall plate anti-rotation structure are exactly the same, and the geometric structure differs only in whether the anti-rotation wall plate is arranged at the leading edge of the blade.

[0040] It should be noted that the geometric plane for parameter comparison shown in the figure is located downstream of the combustion chamber and upstream of the blade. The specific position of this plane can be determined based on the blade leading edge and its chord length. This plane is located upstream of the blade leading edge, and its circumferential distance from the leading edge is 0.5 times the blade axial chord length. This plane has been studied in many literatures as the characteristic surface of the inlet flow field of the turbine first-stage guide vane. The radial distribution of the circumferentially averaged tangential swirl angle and radial swirl angle (averaged along the same radial direction) on this plane shows that the presence of the swirl suppression wall significantly reduces the swirl intensity in the fluid domain. The swirl suppression wall has a stronger effect on the radial swirl angle due to the tangential arrangement of the swirl suppression wall and the orthogonalization of the suppression. Its average swirl intensity is distributed along the normalized blade height, with a suppression of more than 30%. For the tangential swirl angle, the swirl suppression wall has a more significant suppression effect in the upper half, reaching 20%. It can also be seen that for both the radial and tangential swirl angles, the maximum circumferential average inhomogeneity due to structural differences is located near the anti-swirl panel, while the flow field differences at the upper wall away from the anti-swirl panel are not significant. Overall, the anti-swirl panel structure reduces the swirl intensity at the inlet.

[0041] The present invention adopts a wall plate swirl suppression structure diversion method applied to the inlet of the turbine blade, adopts a geometric structure, designs a wall plate swirl suppression structure applied to the inlet of the turbine blade, and places it at a position corresponding to the swirl core. A wall plate swirl suppression structure verification method applied to the inlet of the turbine blade is used to construct a simulation model to verify the structure, and quantitatively analyzes that the swirl suppression structure has a strong inhibitory effect on the upstream swirl flow field, can weaken the uneven swirl in the flow field, increase the uniformity of the flow field, and solves the interference of the combustion chamber outlet swirl on the first-stage guide vane under the strong coupling design of the aircraft engine combustion chamber turbine.

[0042] It should be noted that, in this article, the terms "comprises", "includes" or any other variations thereof are intended to cover non-exclusive inclusion, so that a process, method, article or apparatus that includes a series of elements includes not only those elements, but also includes other elements not explicitly listed, or also includes elements that are inherent to such process, method, article or apparatus.

[0043] The present invention uses specific examples to illustrate the principles and implementation methods of the present invention. The above examples are only used to help understand the method and core ideas of the present invention. The above is only a preferred implementation method of the present invention. It should be pointed out that due to the limitations of textual expression, there are objectively infinite specific structures. For ordinary technicians in this technical field, without departing from the principles of the present invention, they can make several improvements, modifications or changes, and can also combine the above technical features in an appropriate manner; these improvements, modifications, changes or combinations, or the direct application of the inventive concept and technical solution to other occasions without improvement, should be regarded as the scope of protection of the present invention.

Claims

1. A wall plate anti-swirl structure applied to a turbine blade inlet, comprising a fluid domain inlet surface (1), a swirler surface (2), a fluid domain lower wall surface (3), a right blade surface (4), a fluid domain upper wall surface (5), a swirl cylinder channel surface (6), a fluid domain periodic surface (7), an anti-swirl wall plate (8), and a left blade surface (9), characterized in that: The fluid domain inlet surface (1), the cyclone surface (2), the fluid domain lower wall surface (3), the right blade surface (4), the fluid domain upper wall surface (5), the cyclone cylinder channel surface (6), the fluid domain periodic surface (7), and the left blade surface (9) together form a fluid domain unit. The fluid domain periodic surface (7) represents the periodic boundary of the fluid domain unit. The fluid domain unit is generated into the original annular fluid domain after being rotated by the array, wherein the cyclone surface (2) and the cyclone cylinder channel surface (6) are connected to form a cyclone channel. An anti-spin wall plate (8) is provided in the middle of the fluid domain lower wall surface (3) and the fluid domain upper wall surface (5) at the front side of the right blade surface (4) and the left blade surface (9); the anti-spin wall plate (8) is directly connected to the right blade surface (4) and the left blade surface (9) to form a whole. The upstream high-temperature airflow flows through the middle of the flow channel formed by the three.

2. The wall plate anti-rotation structure applied to the turbine blade inlet according to claim 1, characterized in that: The anti-rotation structure comprises an anti-rotation wall plate (8) and a blade surface connected thereto, and corresponds to the leading edge of the first-stage guide vane, and the anti-rotation wall plate (8) is formed into a circular thin plate.

3. The wall plate anti-rotation structure applied to the turbine blade inlet according to claim 1, characterized in that: The anti-rotation wall plate (8) is arranged at the leading edge of the right blade surface (4) and the left blade surface (9) and intersects therewith.

4. The wall plate anti-rotation structure applied to the turbine blade inlet according to claim 1, characterized in that: The angle between the cyclone blades in the cyclone surface (2) and the axis is 45°.

5. A method for using a wall plate anti-rotation structure applied to a turbine blade inlet, comprising the wall plate anti-rotation structure applied to a turbine blade inlet according to claim 1, characterized in that: When the high-pressure gas at the compressor flows into the combustion chamber from the fluid domain inlet surface (1), the gas generates a swirl of a certain intensity after passing through the swirler surface (2) and the swirl cylinder channel surface (6). The swirl suppression wall plate suppresses and controls the swirl, weakens its swirl intensity, and enhances the unevenness of the turbine blade inlet, thereby improving the cooling design effect.

6. A flow guidance verification method for a wall plate anti-spin structure applied to a turbine blade inlet, comprising the wall plate anti-spin structure applied to a turbine blade inlet according to claim 1, characterized by: The steps include: Step 1: Calculate the aerodynamic parameters of the mixed flow field inlet according to the required working conditions; after the temperature is determined, the cyclone is designed according to the working conditions, using the formula , estimate the swirl degree of the designed cyclone; about S in the above formula N Indicates the degree of swirl, the letter D indicates the diameter, where SW is the abbreviation of cyclone, the subscript i indicates the inside, the subscript o indicates the outside, and D SW,i Indicates the inner diameter of the cyclone; Step 2: Select a model with a combustion chamber simulator for modeling and simulation, where the radius of the anti-swirl panel is set at eleven radial positions, evenly distributed between the inner panel radius and the outer panel radius; use RANS to simulate the anti-swirl structure of the panel at different radial positions; after the simulation is completed, the total pressure loss coefficient at the guide vane outlet is used as the evaluation index of aerodynamic loss, and the radial rotation angle and tangential rotation angle are used as indicators of the swirl characteristics of the flow field; after the indicators are selected, the optimal anti-swirl panel placement radius is selected according to the swirl flow field at different swirl center positions, and then a linear relationship is fitted based on the corresponding relationship between the two, which is used as the subsequent empirical formula to facilitate the selection of the anti-swirl panel placement radius; Step 3: Install the swirler on the front side of the unreactive combustion chamber model. Open cooling and mixing holes in the inner and outer cavities of the unreactive combustion chamber. The uniform boundary assumption under multi-porous divergent cooling is used to reduce the amount of data for simulation calculations. Step 4: After being combined with the non-reactive combustion chamber, the high-temperature gas from the upstream combustion chamber flows out from the swirler outlet to form a swirl. After flowing through the anti-swirl wall, the swirl is restricted by the wall, the swirl structure is destroyed, and the swirl of the flow field is weakened.

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