Method for restraining flow separation of surface boundary layer of aero-engine turbine blade

Abstract

The invention discloses a method for restraining flow separation of a boundary layer on the surface of a turbine blade of an aero-engine. The method aims at solving some problems existing in an existing traditional scheme for controlling flow of a turbine of the aero-engine. Flow control is divided into active control and passive control, and a traditional passive control scheme is not flexible enough and has high dependence on flow conditions of a turbine. And an unsteady flowing environment effect is difficult to guarantee. An active control means needs input of external energy and a corresponding control loop, and is complex to realize and high in cost. Aiming at the actual condition that a blade of a high-load low-pressure turbine is objectively vibrated, a certain vibration frequency is directly applied or the original vibration condition is changed so as to avoid a certain frequency or ensure that the vibration frequency of the blade is within a certain safe and flow-beneficial interval; the flow separation on the surface of the turbine blade can be inhibited, so that the turbulence pulsation level in the separation and transition process is reduced or delayed, the total pressure loss is reduced, and the aerodynamic performance of the low-pressure turbine is improved.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft engines, and relates to a method for suppressing boundary layer flow separation on the surface of aircraft engine turbine blades, and specifically to a method for suppressing boundary layer flow separation on the surface of turbine blades based on vibration, which can suppress flow separation of aircraft engine turbine blades and greatly improve the starting performance of aircraft engine turbines. Background Art

[0002] A characteristic of the low-pressure turbine of a civil high-bypass turbofan engine is that it operates in a low Reynolds number range. In the cruising state, the Reynolds number drops significantly due to the sharp decrease in air density. In the aerodynamic environment with a low Reynolds number, a laminar flow state appears in the boundary layer of the suction surface of the low-pressure turbine blade. When passing through the cascade channel, the reverse pressure gradient of the cascade channel increases sharply. The low-speed fluid is not enough to resist the reverse pressure gradient and viscous force, which will cause the boundary layer of the suction surface of the low-pressure turbine blade to easily separate. When the Reynolds number is extremely low, there may even be a large open separation that blocks the flow channel, seriously affecting the aerodynamic performance of the low-pressure turbine.

[0003] In order to improve the performance and reliability of aircraft engine turbine blades, it is necessary to use certain means to suppress or delay the aerodynamic loss caused by flow separation caused by, for example, a low incoming flow Reynolds number. At present, the main control means for suppressing flow separation on the surface of turbine blades in aircraft engines can be divided into two categories from the control mechanism, one is passive control and the other is active control. Passive control refers to affecting the boundary layer state by changing the shape or roughness of the turbine blade wall, such as adding eye sockets, mixing lines, V-grooves and other structures on the blade surface. These control means are simple in structure, easy to implement, and highly reliable. They do not require additional energy to be input into the system. They can induce the transition to a certain extent, thereby delaying or suppressing the boundary layer flow separation and improving the efficiency of the turbine blades. However, while suppressing the separation of airflow, the above-mentioned methods in the prior art have high requirements on the parameters of the passive control structure, are not flexible enough, and are highly dependent on the flow conditions and structural characteristics of the turbine itself. The effect is difficult to guarantee for unsteady flow environments, and there may be adverse effects or instability. , and may affect the cooling effect of the turbine blades. The existing technology also includes active control means to suppress the separation of airflow on the blade surface, which is to change the flow field or boundary conditions by active blowing, wall cooling, etc. to affect the boundary layer state, thereby delaying or suppressing flow separation. The advantage of active control means is that it can be adjusted and optimized according to different working conditions and goals, and has high flexibility and adaptability. However, these methods usually require additional energy input, increase the energy consumption and cost of the system, and require complex control systems and devices, which makes it more difficult to implement to a certain extent, and also increase the structural complexity, failure rate and weight of the aircraft engine turbine.

[0004] In addition, there are some literature (such as "The influence mechanism of low-pressure turbine blade vibration on separation and transition at low Reynolds number", Zhang Yingqiang et al., Journal of Aerospace Power, first published online on December 29, 2022) that uses numerical simulation to compare and analyze the influence of vibrations of low-pressure turbine blades at different frequencies on suction surface boundary layer separation, transition and flow loss at low Reynolds number. The research of these literatures shows that under certain conditions, giving the blades periodic displacement or angle change can effectively suppress flow field separation and improve aerodynamic performance. However, these methods require additional energy input or complex control systems, and may increase blade structural stress and fatigue risk. Therefore, although these methods have certain advantages in theory, they still need to overcome many technical difficulties in practical applications. Therefore, how to use simpler and more effective control methods to suppress flow separation on the surface of turbine blades is one of the problems that need to be solved urgently in current technology. Summary of the invention

[0005] (I) Purpose of the invention

[0006] In view of the above-mentioned defects and shortcomings of the prior art, the present invention aims to provide a method for suppressing boundary layer flow separation on the surface of aircraft engine turbine blades, with the aim of solving some problems existing in the traditional solutions for aircraft engine turbine flow control. Flow control is divided into active control and passive control. The traditional passive control solution is not flexible enough and is highly dependent on the turbine's own flow conditions. The effect is difficult to guarantee for unsteady flow environments. Active control means require the input of external energy and corresponding control loops, which are complex and costly to implement. This brings about problems such as complex structure and heavy weight. , based on the fact that there is no cooling for the low-pressure turbine blades of aircraft engines and the blades of low-pressure turbines are generally thinner, and the blade vibration problem is more likely to occur due to its large aspect ratio, resulting in the objective existence of blade vibration in the low-pressure turbine. A certain vibration frequency is directly applied to the blades through an excitation device or an excitation or damping device is added to change the original vibration conditions, and even the mass and rigidity of the blades can be changed to change the natural frequency of the blades, so as to avoid a certain frequency or ensure that the blade vibration frequency is within a safe and flow-friendly range, which can inhibit the flow separation on the surface of the turbine blades, reduce or delay the level of turbulent pulsation during separation and transition, and reduce the total pressure loss, thereby improving and improving the aerodynamic performance of the low-pressure turbine. At the same time, compared with the prior art, the present invention does not require or input a small amount of additional energy, does not require the addition of structures to cause additional aerodynamic losses, does not require the modification of the blades themselves, does not require the addition of complex structures to the turbine components of aircraft engines, and has better flexibility and adaptability to the internal structure and flow field conditions of the turbine.

[0007] (II) Technical solution

[0008] In order to achieve the invention purpose and solve the technical problem, the technical solution adopted by the present invention is as follows:

[0009] A method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade, characterized in that the suppression method at least comprises the following steps:

[0010] SS1.According to the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine, the working pressure of the low-pressure turbine and the working Reynolds number Re representing the turbulence degree of the incoming air flow are obtained;

[0011] SS2. Based on the working pressure and working Reynolds number of the low-pressure turbine calculated in step SS1, a CFD numerical simulation model of the low-pressure turbine in a high-altitude environment is established, and on the basis of the CFD numerical simulation model, multiple groups of different vibration modes, vibration amplitudes and equivalent vibration frequencies are set to fully simulate the vibration characteristics of the low-pressure turbine under aerodynamic conditions, wherein the vibration mode is the deformation mode of the low-pressure turbine blades at different frequencies, and the equivalent vibration frequency is the actual vibration frequency of the low-pressure turbine blades after considering the influence of airflow on the vibration;

[0012] SS3. Based on the numerical simulation model of the low-pressure turbine in a high-altitude environment established in step SS2, without considering the influence of vibration, the flow field of the low-pressure turbine blade in a high-altitude environment is numerically simulated by numerical simulation and supplemented by theoretical analysis to obtain the two-dimensional flow characteristics and important aerodynamic performance parameters of the low-pressure turbine blade in a high-altitude environment, wherein the important aerodynamic performance parameters at least include the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade, and the two-dimensional flow characteristics at least include the velocity distribution, pressure distribution and / or temperature distribution of the airflow in the plane of the low-pressure turbine blade, the separation bubble size is used to characterize the size of the laminar separation area formed by the airflow on the surface of the low-pressure turbine blade due to the adverse pressure gradient, and the total pressure loss is used to characterize the difference between the total pressure after the airflow passes through the low-pressure turbine blade and the inlet total pressure, so as to reflect the efficiency of the low-pressure turbine blade in doing work on the airflow;

[0013] SS4. Based on the numerical simulation model of the low-pressure turbine in the high-altitude environment established in step SS2, and based on the set vibration mode, vibration amplitude and equivalent vibration frequency, under the premise of considering the influence of vibration, through numerical simulation and supplemented by theoretical analysis, the flow field change caused by the vibration excitation of the low-pressure turbine blade is obtained, the influence of vibration on flow separation is obtained, and the relationship between the separation bubble size of the boundary layer on the blade surface and the total pressure loss of the low-pressure turbine blade and the equivalent vibration frequency of the blade is established;

[0014] SS5. Based on the relationship between the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade and the blade equivalent vibration frequency established in step SS4 and the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade when there is no vibration influence obtained in step SS3, a comparative analysis is performed between the two to obtain the separation bubble size of the blade surface boundary layer of the low-pressure turbine blade in a high-altitude environment and the optimal blade equivalent vibration frequency when the total pressure loss of the low-pressure turbine blade is minimized;

[0015] SS6. According to the optimal blade reduced vibration frequency corresponding to the minimum separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade in a high-altitude environment obtained after comparative analysis in step SS5, the low-pressure turbine blade is subjected to vibration by excitation or the main vibration frequency is controlled to be near the optimal blade reduced vibration frequency by other means to minimize the blade profile loss of the low-pressure turbine blade in a high-altitude environment.

[0016] Preferably, in the above step SS1, the operating Reynolds number Re of the low-pressure turbine is obtained according to the following steps:

[0017] SS1.1 Determine the temperature and pressure of the incoming air flow according to the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine, and calculate the density of the incoming air flow according to the ideal gas state equation:

[0018]

[0019] Among them, ρ is the density of the air flow, T is the temperature of the air flow, p is the pressure of the air flow, and R is the gas constant of the air;

[0020] SS1.2 Calculate the dynamic viscosity μ of the incoming air flow based on the temperature T of the incoming air flow and the empirical formula for calculating the dynamic viscosity:

[0021]

[0022] Among them, μ0 and T0 are the dynamic viscosity and temperature of air under the reference state;

[0023] SS1.3 Determine the air velocity V and the characteristic length D of the low-pressure turbine blades based on the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine;

[0024] SS1.4 obtains the working Reynolds number Re of the low-pressure turbine according to the following formula:

[0025]

[0026] Among them, ρ is the density of the incoming air flow, V is the velocity of the incoming air flow, D is the characteristic length of the low-pressure turbine blade, and μ is the dynamic viscosity of the incoming air flow.

[0027] Preferably, in the above step SS2, the CFD numerical simulation model is solved based on the Reynolds-averaged Navier-Stokes equations, and the turbulence simulation of the boundary layer on the surface of the low-pressure turbine blade is performed using the k-ε turbulence model, k-ω turbulence model, LES model, SST turbulence model or DNS model, and the γ-Re transition model is used to predict the transition position of the boundary layer flow.

[0028] Preferably, in the above step SS2, the vibration mode includes at least translational vibration, rotational vibration or bending vibration, and a suitable vibration mode is selected based on the working characteristics of the low-pressure turbine blade in a high-altitude environment. These vibration modes can respectively simulate the deformation of the blade in different directions, and take into account the geometric nonlinearity and aerodynamic nonlinearity effects of the blade.

[0029] Preferably, in the above step SS2, the range of the reduced vibration frequency is 0 to 17.5. More preferably, the range of the reduced vibration frequency is 0 to 15. The reason for this setting is that, through numerical calculations, it is found that when the vibration frequency of the low-pressure turbine blade exceeds 15, the separation bubble profile of the boundary layer on the blade surface will change, and secondary separation bubbles will also be generated. Secondary separation bubbles will increase blade profile losses and should be avoided as much as possible.

[0030] Preferably, in the above step SS2, the vibration amplitude is set to about 1.5% of the chord length of the low-pressure turbine blade. When setting the vibration amplitude, if the vibration amplitude is too small, it will lead to insufficient excitation effect, making it difficult to effectively destroy the flow separation of the boundary layer on the blade surface, and it will be difficult to achieve the purpose of suppressing separation. In addition, if the vibration amplitude is too large, it will have a greater impact on the structural strength of the turbine blades, easily leading to fatigue failure and affecting the service life of the turbine. 1.5% of the blade chord length is a relatively mild vibration amplitude value. Considering the accuracy of the vibration excitation device in engineering applications, this vibration amplitude is easier to achieve and control, and is more conducive to engineering applications.

[0031] Preferably, in the above step SS4, when establishing the relationship between the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade and the blade equivalent vibration frequency, a data fitting technology based on an empirical formula or a machine learning method such as a neural network is used. This technology can improve the fitting accuracy and efficiency and can handle complex nonlinear relationships.

[0032] Preferably, in the above step SS6, the excitation method at least includes electromagnetic excitation, mechanical excitation, fluidic excitation and / or piezoelectric excitation.

[0033] Furthermore, in the above step SS6, the vibration excitation is set on the suction surface of the low-pressure turbine blade and is set at the middle and rear chord length position near the trailing edge of the blade. The reason is that boundary layer flow separation usually occurs near the trailing edge, while the flow characteristics from the middle of the blade to the leading edge are less affected by the separation.

[0034] (III) Technical Effect

[0035] Compared with the prior art, the method for suppressing boundary layer flow separation on the surface of aeroengine turbine blades proposed by the present invention has significant technical effects:

[0036] 1. Since there is no cooling problem for low-pressure turbine blades, the blades are generally thinner, and due to their large aspect ratio, blade vibration problems are more likely to occur. The actual low-pressure turbine does have blade vibration problems. The present invention basically does not require additional energy input, does not require modification of the blades themselves, and does not require the addition of complex structures to the aircraft engine turbine components. It only requires adding a simple vibration excitation device to the existing low-pressure turbine and starting vibration excitation under certain working conditions (when the inherent vibration of the low-pressure turbine blades themselves meets the corresponding conditions, there is no need to start vibration excitation).

[0037] 2. The method for suppressing the flow separation of the boundary layer on the surface of the turbine blade of an aircraft engine proposed in the present invention establishes the relationship between the separation bubble size of the boundary layer on the surface of the blade and the total pressure loss of the low-pressure turbine blade and the blade reduced vibration frequency through numerical simulation and theoretical analysis, and determines the optimal blade reduced vibration frequency based on the relationship, thereby optimizing the flow control;

[0038] 3. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade proposed in the present invention applies appropriate vibration excitation to the low-pressure turbine blade to change the flow state on the blade surface, suppress boundary layer flow separation, reduce blade profile loss, and thus improve the efficiency of the low-pressure turbine in doing work on the airflow, and improve the fuel economy and thrust of the aircraft engine;

[0039] 4. The method for suppressing boundary layer flow separation on the surface of aircraft engine turbine blades proposed in the present invention can flexibly adjust the parameters of vibration excitation according to different aircraft engine models and different flight altitudes and environmental conditions to adapt to different flow field requirements. At the same time, a variety of excitation methods can be adopted, such as electromagnetic excitation, mechanical excitation or jet excitation, to minimize the blade loss of low-pressure turbine blades in high-altitude environments. It can effectively improve the efficiency and stability of the turbine under different working conditions, realize the optimal control of complex three-dimensional flow fields, and has important engineering application value. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1The present invention is a schematic diagram of the working process of the method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade.

[0041] Figure 2 Schematic diagram of the relationship between the separation bubble size and the blade reduced vibration frequency.

[0042] Figure 3 Schematic diagram of the relationship between total pressure loss and blade reduced vibration frequency. DETAILED DESCRIPTION

[0043] In order to better understand the present invention, the content of the present invention is further explained below in conjunction with the embodiments, so that the advantages and features of the present invention can be more easily understood by those skilled in the art. It should be noted that the following is only a preferred embodiment of the present invention, but the content of the present invention is not limited to the following embodiments. In fact, various modifications and changes can be made in the present invention without departing from the scope or spirit of the present invention, which will be apparent to those skilled in the art. For example, the features shown or described as a part of an embodiment can be used together with another embodiment to produce another embodiment. Therefore, it is intended that the present invention includes such modifications and changes within the scope of the appended claims and their equivalents.

[0044] like Figure 1 As shown, the method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade of the present invention comprises at least the following steps when implemented:

[0045] SS1.According to the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine, the working pressure of the low-pressure turbine and the working Reynolds number Re representing the turbulence degree of the incoming air flow are obtained;

[0046] SS2. Based on the working pressure and working Reynolds number of the low-pressure turbine calculated in step SS1, a CFD numerical simulation model of the low-pressure turbine in a high-altitude environment is established, and on the basis of the CFD numerical simulation model, multiple groups of different vibration modes, vibration amplitudes and equivalent vibration frequencies are set to fully simulate the vibration characteristics of the low-pressure turbine under aerodynamic conditions, wherein the vibration mode is the deformation mode of the low-pressure turbine blades at different frequencies, and the equivalent vibration frequency is the actual vibration frequency of the low-pressure turbine blades after considering the influence of airflow on the vibration;

[0047] SS3. Based on the numerical simulation model of the low-pressure turbine in a high-altitude environment established in step SS2, without considering the influence of vibration, the flow field of the low-pressure turbine blade in a high-altitude environment is numerically simulated by numerical simulation and supplemented by theoretical analysis to obtain the two-dimensional flow characteristics and important aerodynamic performance parameters of the low-pressure turbine blade in a high-altitude environment, wherein the important aerodynamic performance parameters at least include the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade, and the two-dimensional flow characteristics at least include the velocity distribution, pressure distribution and / or temperature distribution of the airflow in the plane of the low-pressure turbine blade, the separation bubble size is used to characterize the size of the laminar separation area formed by the airflow on the surface of the low-pressure turbine blade due to the adverse pressure gradient, and the total pressure loss is used to characterize the difference between the total pressure after the airflow passes through the low-pressure turbine blade and the inlet total pressure, so as to reflect the efficiency of the low-pressure turbine blade in doing work on the airflow;

[0048] SS4. Based on the numerical simulation model of the low-pressure turbine in the high-altitude environment established in step SS2, and based on the set vibration mode, vibration amplitude and equivalent vibration frequency, under the premise of considering the influence of vibration, through numerical simulation and supplemented by theoretical analysis, the flow field change caused by the vibration excitation of the low-pressure turbine blade is obtained, the influence of vibration on flow separation is obtained, and the relationship between the separation bubble size of the boundary layer on the blade surface and the total pressure loss of the low-pressure turbine blade and the equivalent vibration frequency of the blade is established;

[0049] SS5. Based on the relationship between the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade and the blade equivalent vibration frequency established in step SS4 and the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade when there is no vibration influence obtained in step SS3, a comparative analysis is performed between the two to obtain the separation bubble size of the blade surface boundary layer of the low-pressure turbine blade in a high-altitude environment and the optimal blade equivalent vibration frequency when the total pressure loss of the low-pressure turbine blade is minimized;

[0050] SS6. According to the optimal blade reduced vibration frequency corresponding to the minimum separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade in a high-altitude environment obtained after comparative analysis in step SS5, the low-pressure turbine blade is subjected to vibration by excitation or the main vibration frequency is controlled to be near the optimal blade reduced vibration frequency by other means to minimize the blade profile loss of the low-pressure turbine blade in a high-altitude environment.

[0051] In summary, the method of suppressing the flow separation of the boundary layer on the surface of the turbine blade of an aeroengine of the present invention has the following working principle: by establishing a CFD numerical simulation model of a low-pressure turbine in a high-altitude environment, and setting multiple groups of different vibration modes, vibration amplitudes and equivalent vibration frequencies, the vibration characteristics of the low-pressure turbine under aerodynamic conditions are fully simulated. On this basis, by numerical simulation and supplemented by theoretical analysis, the flow field changes caused by the vibration excitation of the low-pressure turbine blade are obtained, the influence of vibration on the flow separation is obtained, and the relationship between the separation bubble size of the boundary layer on the blade surface and the total pressure loss of the low-pressure turbine blade and the blade equivalent vibration frequency is established. Finally, by comparative analysis, the separation bubble size of the boundary layer on the blade surface of the low-pressure turbine blade in a high-altitude environment and the optimal blade equivalent vibration frequency when the total pressure loss of the low-pressure turbine blade is minimized are obtained, and the vibration corresponding to the optimal blade equivalent vibration frequency is applied to the low-pressure turbine blade by excitation to minimize the blade profile loss of the low-pressure turbine blade in a high-altitude environment. The method can effectively suppress the flow separation of the boundary layer on the surface of the turbine blade of an aeroengine and improve the efficiency and reliability of the engine.

[0052] More specifically, corresponding to the above step SS1, according to a certain cruising altitude of the aircraft engine and the working environment of the low-pressure turbine, the working Reynolds number of the low-pressure turbine is obtained as Re=25000.

[0053] Corresponding to the above step SS2, the numerical simulation parameters of the low-pressure turbine working pressure and Reynolds number simulating the high-altitude environment are calculated, and the vibration mode is selected as translation, and the equivalent vibration frequency range is 0 to 17.5. The equivalent vibration frequency K is a dimensionless parameter, and its calculation formula is Where C is the axial chord length of the low-pressure turbine blade, V ref is the reference speed of the incoming air flow, and T is the vibration period of the low-pressure turbine blades.

[0054] Corresponding to the above step SS4, based on the numerical simulation model of the low-pressure turbine in the high-altitude environment established in step SS2, and based on the set vibration mode, vibration amplitude and equivalent vibration frequency, under the premise of considering the influence of vibration, the flow field of the low-pressure turbine blade is simulated through theoretical analysis and numerical simulation to obtain the aerodynamic performance parameters such as the separation bubble size and total pressure loss of the low-pressure turbine blade, such as Figure 2 , Figure 3 shown.

[0055] Figure 2 , Figure 3 It shows that when the blade reduced vibration frequency is below 15, the separation bubble size will continue to decrease, and the total pressure loss coefficient will decrease with the increase of the blade reduced vibration frequency. When the blade reduced vibration frequency exceeds 15 and continues to increase, the total pressure loss will increase slightly. Figure 3The relationship between the total pressure loss and the blade reduced vibration frequency is shown in the figure. Electromagnetic excitation, mechanical excitation, jet excitation and piezoelectric excitation are used to apply vibration with a reduced frequency below 15 to the blade to achieve the purpose of reducing blade loss.

[0056] In a preferred embodiment of the present invention, in the above step SS1, the working Reynolds number Re of the low-pressure turbine is obtained according to the following steps:

[0057] SS1.1 Determine the temperature and pressure of the incoming air flow according to the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine, and calculate the density of the incoming air flow according to the ideal gas state equation:

[0058]

[0059] Among them, ρ is the density of the air flow, T is the temperature of the air flow, p is the pressure of the air flow, and R is the gas constant of the air;

[0060] SS1.2 Calculate the dynamic viscosity μ of the incoming air flow based on the temperature T of the incoming air flow and the empirical formula for calculating the dynamic viscosity:

[0061]

[0062] Among them, μ0 and T0 are the dynamic viscosity and temperature of air under the reference state;

[0063] SS1.3 Determine the air velocity V and the characteristic length D of the low-pressure turbine blades based on the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine;

[0064] SS1.4 obtains the working Reynolds number Re of the low-pressure turbine according to the following formula:

[0065]

[0066] Among them, ρ is the density of the incoming air flow, V is the velocity of the incoming air flow, D is the characteristic length of the low-pressure turbine blade, and μ is the dynamic viscosity of the incoming air flow.

[0067] In a preferred embodiment of the present invention, in the above step SS2, the CFD numerical simulation model is solved based on the Reynolds-averaged Navier-Stokes equations, and the k-ε turbulence model, k-ω turbulence model, LES model, SST turbulence model or DNS model are used to simulate the turbulence of the boundary layer on the surface of the low-pressure turbine blade, and the γ-Re transition model is used to predict the transition position of the boundary layer flow. The vibration mode includes at least translational vibration, rotational vibration or bending vibration, and a suitable vibration mode is selected based on the working characteristics of the low-pressure turbine blade in a high-altitude environment. These vibration modes can respectively simulate the deformation of the blade in different directions, and consider the geometric nonlinearity and aerodynamic nonlinearity of the blade. The set vibration amplitude is about 1.5% of the chord length of the low-pressure turbine blade. When setting the vibration amplitude, if the vibration amplitude is too small, it will lead to insufficient excitation effect, it is difficult to effectively destroy the flow separation of the boundary layer on the surface of the blade, and it is difficult to achieve the purpose of suppressing separation. In addition, if the vibration amplitude is too large, it will have a greater impact on the structural strength of the turbine blade, easily lead to fatigue failure, and affect the service life of the low-pressure turbine blade. 1.5% of the blade chord length is a relatively mild vibration amplitude value. Considering the accuracy of the vibration excitation device in engineering applications, this vibration amplitude is easier to achieve and control, and is more conducive to engineering applications.

[0068] In a preferred embodiment of the present invention, in the above step SS4, when establishing the relationship between the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade and the blade equivalent vibration frequency, a data fitting technology based on an empirical formula or a machine learning method such as a neural network is used. This technology can improve the fitting accuracy and efficiency, and can handle complex nonlinear relationships.

[0069] Through the above embodiments, the purpose of the present invention is fully and effectively achieved. Any equivalent or simple changes made according to the structure, features and principles described in the patent concept of the present invention are included in the protection scope of the patent of the present invention. The technicians in the technical field of the present invention can make various modifications or supplements to the specific embodiments described or replace them in a similar manner, as long as they do not deviate from the structure of the present invention or exceed the scope defined by the claims, they should all belong to the protection scope of the present invention.

Claims

1. A method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade, characterized in that: The inhibition method at least comprises the following steps: SS1.According to the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine, the working pressure of the low-pressure turbine and the working Reynolds number Re representing the turbulence degree of the incoming air flow are obtained; SS2. Based on the working pressure and working Reynolds number of the low-pressure turbine calculated in step SS1, a CFD numerical simulation model of the low-pressure turbine in a high-altitude environment is established, and on the basis of the CFD numerical simulation model, multiple groups of different vibration modes, vibration amplitudes and equivalent vibration frequencies are set to fully simulate the vibration characteristics of the low-pressure turbine under aerodynamic conditions, wherein the vibration mode is the deformation mode of the low-pressure turbine blades at different frequencies, and the equivalent vibration frequency is the actual vibration frequency of the low-pressure turbine blades after considering the influence of airflow on the vibration; SS3. Based on the numerical simulation model of the low-pressure turbine in a high-altitude environment established in step SS2, without considering the influence of vibration, the flow field of the low-pressure turbine blade in a high-altitude environment is numerically simulated by numerical simulation and supplemented by theoretical analysis to obtain the two-dimensional flow characteristics and important aerodynamic performance parameters of the low-pressure turbine blade in a high-altitude environment, wherein the important aerodynamic performance parameters at least include the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade, and the two-dimensional flow characteristics at least include the velocity distribution, pressure distribution and / or temperature distribution of the airflow in the plane of the low-pressure turbine blade, the separation bubble size is used to characterize the size of the laminar separation area formed by the airflow on the surface of the low-pressure turbine blade due to the adverse pressure gradient, and the total pressure loss is used to characterize the difference between the total pressure after the airflow passes through the low-pressure turbine blade and the inlet total pressure, so as to reflect the efficiency of the low-pressure turbine blade in doing work on the airflow; SS4. Based on the numerical simulation model of the low-pressure turbine in the high-altitude environment established in step SS2, and based on the set vibration mode, vibration amplitude and equivalent vibration frequency, under the premise of considering the influence of vibration, through numerical simulation and supplemented by theoretical analysis, the flow field change caused by the vibration excitation of the low-pressure turbine blade is obtained, the influence of vibration on flow separation is obtained, and the relationship between the separation bubble size of the boundary layer on the blade surface and the total pressure loss of the low-pressure turbine blade and the equivalent vibration frequency of the blade is established; SS5. Based on the relationship between the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade and the blade equivalent vibration frequency established in step SS4 and the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade when there is no vibration influence obtained in step SS3, a comparative analysis is performed between the two to obtain the separation bubble size of the blade surface boundary layer of the low-pressure turbine blade in a high-altitude environment and the optimal blade equivalent vibration frequency when the total pressure loss of the low-pressure turbine blade is minimized; SS6. According to the optimal blade reduced vibration frequency corresponding to the minimum separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade in a high-altitude environment obtained after comparative analysis in step SS5, the low-pressure turbine blade is subjected to vibration by excitation or the main vibration frequency is controlled to be near the optimal blade reduced vibration frequency by other means to minimize the blade profile loss of the low-pressure turbine blade in a high-altitude environment.

2. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 1, characterized in that: In the above step SS1, the working Reynolds number Re of the low-pressure turbine is obtained according to the following steps: SS1.1 Determine the temperature and pressure of the incoming air flow according to the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine, and calculate the density of the incoming air flow according to the ideal gas state equation: Among them, ρ is the density of the air flow, T is the temperature of the air flow, p is the pressure of the air flow, and R is the gas constant of the air; SS1.2 Calculate the dynamic viscosity μ of the incoming air flow based on the temperature T of the incoming air flow and the empirical formula for calculating the dynamic viscosity: Among them, μ0 and T0 are the dynamic viscosity and temperature of air under the reference state; SS1.3 Determine the air velocity V and the characteristic length D of the low-pressure turbine blades based on the cruising altitude of the aircraft engine and the working environment of the low-pressure turbine; SS1.4 obtains the working Reynolds number Re of the low-pressure turbine according to the following formula: Among them, ρ is the density of the incoming air flow, V is the velocity of the incoming air flow, D is the characteristic length of the low-pressure turbine blade, and μ is the dynamic viscosity of the incoming air flow.

3. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 1, characterized in that: In the above step SS2, the CFD numerical simulation model is solved based on the Reynolds-averaged Navier-Stokes equations, and the k-ε turbulence model, k-ω turbulence model, LES model, SST turbulence model or DNS model are used to simulate the turbulence of the boundary layer on the surface of the low-pressure turbine blade, and the γ-Re transition model is used to predict the transition position of the boundary layer flow.

4. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 1, characterized in that: In the above step SS2, the vibration mode includes at least translational vibration, rotational vibration or bending vibration, and a suitable vibration mode is selected based on the working characteristics of the low-pressure turbine blades in a high-altitude environment.

5. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 1, characterized in that: In the above step SS2, the range of the converted vibration frequency is 0 to 17.

5.

6. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 5, characterized in that: In the above step SS2, the range of the converted vibration frequency is 0-15.

7. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 1, characterized in that: In the above step SS2, the vibration amplitude is set to be about 1.5% of the chord length of the low-pressure turbine blade.

8. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 1, characterized in that: In the above step SS4, when establishing the relationship between the separation bubble size of the blade surface boundary layer and the total pressure loss of the low-pressure turbine blade and the blade reduced vibration frequency, a data fitting technology based on empirical formulas or machine learning methods such as neural networks is used.

9. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 1, characterized in that: In the above step SS6, the excitation method at least includes electromagnetic excitation, mechanical excitation, fluidic excitation and / or piezoelectric excitation.

10. The method for suppressing boundary layer flow separation on the surface of an aircraft engine turbine blade according to claim 8, characterized in that: In the above step SS6, the vibration excitation is set on the suction surface of the low-pressure turbine blade and is set at the middle and rear chord length position close to the trailing edge of the blade.

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