Self-adaptive cooling structure for turbine guide vane of aero-engine

Abstract

The invention discloses a self-adaptive cooling structure of an aero-engine turbine guide vane, and belongs to the field of turbine guide vane design. The aero-engine turbine guide vane comprises a vane body, and at least one cooling cavity is formed in the vane body. A self-adaptive adjusting unit is arranged on the cooling cavity, and the self-adaptive adjusting unit dynamically adjusts the flow or the flow direction of a cooling medium flowing through the cooling cavity based on the thermal load change of the working condition where the turbine guide vane is located, so that self-adaptive control over the cooling strength of the vane body is achieved. The structure has the self-adaptive characteristic, namely, the cooling airflow distribution or the structural form can be dynamically adjusted according to working conditions (such as temperature, pressure and rotating speed).

Description

Technical Field

[0001] This invention belongs to the field of turbine guide vane design, specifically relating to an adaptive cooling structure for aero-engine turbine guide vanes. Background Technology

[0002] With the continuous improvement of aero-engine performance, turbine inlet temperatures have increased significantly, placing extremely high demands on the cooling efficiency of turbine guide vanes. Impact cooling is one of the key internal cooling technologies for modern high-efficiency air-cooled blades. However, the impact cooling structures in existing technologies (such as the size and location of impact holes / slots) are fixed. This fixed structure is optimized based on a specific design point and cannot adapt to the complex operating conditions of engines in actual operation.

[0003] When an aircraft flies at different altitudes and speeds, the parameters (such as pressure and temperature) at the guide vane inlet undergo significant changes. Fixed-structure cooling systems deviate significantly from their optimal performance outside of design points, especially under high-temperature and high-pressure conditions. This can lead to problems such as improper cooling air distribution and weakened impact effects, resulting in insufficient blade cooling efficiency and localized overheating, seriously threatening engine safety and reliability. Existing solutions for improving blade cooling efficiency typically take the following forms: 1. A more complex multi-channel impact cooling structure is used to achieve efficient cooling of the blades under high operating conditions; 2. Use irregularly shaped holes to improve cooling efficiency; 3. Multiple fins are added to enhance internal convection heat transfer; All of the aforementioned existing technical solutions cannot overcome the inherent defects caused by their "fixed geometry" characteristics (e.g., fixed-diameter impact holes or impact slit arrays). Under high operating conditions, their fixed flow area leads to insufficient cooling air volume and weakened impact effect, easily causing blade overheating. Existing turbine blade impact cooling structures suffer from the following common and fatal drawbacks due to their fixed geometry: 1. Inability to adapt to changes in operating conditions (core defect): It cannot respond to changes in engine operating conditions (guide vane inlet pressure, temperature) caused by changes in aircraft altitude and speed. Its design can only be optimized based on a specific operating condition (such as cruise state), and once it deviates from that condition, the cooling performance deviates sharply from the optimal value.

[0004] 2. Insufficient cooling capacity under high operating conditions: The growth rate of external heat load on the blades far exceeds the cooling capacity of the fixed cooling structure, resulting in "overheating" of the blades and a relative decrease in cooling efficiency.

[0005] 3. High manufacturing and maintenance costs: Complex air conditioning channels require high-precision casting technology and expensive special processing (laser drilling), resulting in high manufacturing costs.

[0006] Therefore, there is an urgent need in this field for a structure that can intelligently adapt to changes in flight conditions and maintain efficient cooling throughout the entire flight envelope, in order to solve the aforementioned long-standing technical challenges. Summary of the Invention

[0007] The technical problem to be solved: To avoid the shortcomings of existing technologies, this invention provides an adaptive cooling structure for turbine guide vanes of aero-engines, which has an "adaptive" characteristic, that is, it can dynamically adjust the cooling airflow distribution or structural shape according to operating conditions (such as temperature, pressure, and speed).

[0008] The technical solution of the present invention is: an adaptive cooling structure for aero-engine turbine guide vanes, wherein the aero-engine turbine guide vane includes a blade body, and at least one cooling cavity is provided inside the blade body; an adaptive adjustment unit is provided on the cooling cavity, and the adaptive adjustment unit dynamically adjusts the flow rate or direction of the cooling medium flowing through the cooling cavity based on the thermal load change of the turbine guide vane under the operating conditions, so as to achieve adaptive control of the cooling intensity of the blade body.

[0009] A further technical solution of the present invention is: the adaptive adjustment unit includes a flow adjustment component disposed in the circumference of at least one cooling cavity, the flow adjustment component being able to change its own opening degree or flow area in response to temperature or pressure changes in the area surrounding the cooling cavity. A further technical solution of the present invention is: the cooling cavity has multiple impact slits extending through the span of the blades along the circumferential direction, and the lip of the impact slit is composed of a flow regulating component; the flow regulating component is a shape memory alloy actuator, and when the local temperature rises to the phase transition trigger temperature of the shape memory alloy, the shape memory alloy actuator deforms, thereby increasing the effective flow area of ​​the impact slit.

[0010] A further technical solution of the present invention is: the shape memory alloy actuator is a bimetallic strip, which is respectively formed or connected to the walls on both sides of the impact slit; the bimetallic strip undergoes bending deformation based on temperature changes, thereby adjusting the opening of the impact slit.

[0011] A further technical solution of the present invention is that the bimetallic sheet is made of gold.

[0012] A further technical solution of the present invention is: the cooling cavity inside the blade includes a front cavity and a rear cavity; The wall of the front cavity is provided with five impact slits along the circumferential direction, and the widths of the bimetallic strips are: SH1=0.53mm, SH2=0.65mm, SH3=0.518984mm, SH4=0.518984mm, SH5=0.659680mm respectively. The rear cavity has five impact slits along its circumferential direction on its wall surface. Two slits are provided on the pressure side, with bimetallic strip widths of PS1=0.44267mm and PS2=0.45700mm respectively; and three slits are provided on the suction side, with bimetallic strip widths of SS1=0.27829mm, SS2=0.30289mm and SS3=0.29890mm respectively.

[0013] A design method for an adaptive cooling structure for aero-engine turbine guide vanes includes the following steps: Step 1: Define the effective uniformly distributed load ; The effective uniformly distributed load To represent the ideal uniformly distributed load equivalent to the complex pressure distribution generated by the jet airflow from the impact slit, it is determined by the layout P, ​​number N, width W, and total pressure at the main inlet of the impact slit. and the total temperature of the main entrance The function is represented as:

[0014] Step 2: Calculate the deformation of the metal sheet in the direction perpendicular to it. y z ; According to the effective uniformly distributed load The deformation was calculated using the beam bending deformation formula. y z :

[0015] in, l denoted as , where E is the length of the metal sheet; E is the elastic modulus of the metal sheet; and I is the moment of inertia of the cross section. Step 3: Calculate the deformation of the metal sheet in the direction parallel to it. y x ; Considering the distributed torque T generated simultaneously by the gas impact, first calculate the torsion angle at the end of the support plate. :

[0016] in, G It is the shear modulus of the material. J It is the torsional moment of inertia of the cross section; Subsequently, based on the stated torsion angle and vertical deformation y x Calculate the deformation in the parallel direction. :

[0017] Step 4: Establish the relationship between deformation and changes in slit cross-sectional area to achieve adaptive flow rate adjustment; Based on the calculated deformation in the parallel direction Determine the change in the effective flow cross-sectional area of ​​the impact slit, wherein the change in cross-sectional area is related to... Positive correlation; by iteratively optimizing the layout P, ​​number N, and width W parameters of the impact slits, the cross-sectional area variation can adapt to the total pressure at the main inlet. and the total temperature of the main entrance The changes in temperature allow for self-regulation of the cooling flow rate.

[0018] A parametric optimization design method for an adaptive cooling structure of aero-engine turbine guide vanes, comprising the following steps: Step 1: Establish analysis baseline Based on the three-dimensional model of the turbine guide vane including the initial impact slit layout, flow field simulation analysis is performed to obtain the baseline heat load distribution and cooling efficiency distribution on the blade surface. Step 2: Set optimization objectives and constraints The optimization objective is to improve overall cooling efficiency, and the core constraint is that the maximum surface temperature of the blades is lower than the allowable temperature of the material. Step 3: Execute iterative optimization loop Repeat the following sub-steps until the optimization objective and constraints set in step 2 are met: Step 3.1: Simulation and Analysis: Perform flow field simulation on the current version of the turbine guide vane model to obtain its temperature field and cooling efficiency data; Step 3.2: Strategy Formulation: Compare the current simulation results with the baseline heat load distribution to identify areas with insufficient cooling and areas with excessive cooling, and formulate adjustment strategies for the layout, number, and / or width of the impact slits based on this. Step 3.3: Model Update: Based on the adjustment strategy, the geometric parameters of the impact slit in the turbine guide vane model are modified parametrically to generate a new model for the next iteration.

[0019] An aero-engine, wherein the high-pressure turbine guide vane section adopts the adaptive cooling structure of the aero-engine turbine guide vane.

[0020] Beneficial effects The beneficial effects of this invention are as follows: This invention designs an adaptive cooling structure for turbine guide vanes, which can dynamically and continuously adjust the flow area of ​​the cooling channel in response to changes in pressure parameters through an elastic deformable wall. Compared with existing fixed geometry technology, this invention breaks through the limitation of traditional designs being optimal only under a single operating condition. Under high operating conditions, the slit automatically expands due to increased pressure, increasing the cooling flow rate and effectively solving the risk of overheating of the blades due to insufficient cooling; while under low operating conditions, the slit can automatically maintain a relatively narrow state to limit the flow rate, realizing on-demand distribution of cooling air volume, thereby improving the overall efficiency of the engine across the entire operating range. Specific effects are as follows: 1. This invention introduces an adaptive adjustment unit, enabling the cooling system to sense and respond to changes in the engine's actual operating state. Under high heat load conditions, it automatically enhances cooling capacity to prevent blade overheating; under low heat load conditions, it reduces cooling air consumption. This overcomes the limitation of traditional fixed structures being optimal only at the design point, achieving dynamic optimization of cooling efficiency throughout the entire flight envelope.

[0021] 2. By avoiding excessive use of cooling air under off-design conditions, this invention effectively reduces the amount of cooling air drawn from the compressor. This saved high-pressure air can be used more for combustion and power generation, directly improving engine thrust and thermal efficiency, which is of great significance for reducing fuel consumption and operating costs.

[0022] In summary, by upgrading the cooling system of the turbine guide vane from "static design" to "dynamic adaptive design", this invention effectively balances the two core and contradictory requirements of blade safety and engine efficiency, providing key technical support for the development of the next generation of high-performance aero engines. Attached Figure Description

[0023] Figure 1 Adaptive impact slit structure; (a) front cavity impact slit, (b) rear cavity impact slit; Figure 2 : Structure before and after deformation of the impact slit; (a) Mid-section structure before deformation of the impact slit, (b) Mid-section structure after deformation of the impact slit; Figure 3 Comparison of structural cloud maps before and after impact slit deformation; (a) Structural cooling effect cloud map before impact slit deformation, with an average comprehensive cooling effect of 0.522547; (b) Structural cooling effect cloud map after impact slit deformation, with an average comprehensive cooling effect of 0.551006. Figure 4 The overall cooling effect distribution of the cross section of the guide vane before and after deformation; Figure 5 Optimization Design Method for Adaptive Cooling Structure of Turbine Guide Blades Explanation of reference numerals in the attached diagram: SH1-SH5 are five impact slots in the front cavity of the blade; PS1 and PS2 are two impact slots on the pressure side of the rear cavity of the blade; SS1-SS3 are three impact slots on the suction side of the rear cavity of the blade. Detailed Implementation

[0024] The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain the invention, and should not be construed as limiting the invention.

[0025] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," and "counterclockwise," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0026] In the existing technology, the cooling structure design of turbine guide vanes has developed a variety of mature solutions. For example, multi-chamber structures (such as CN201811230895.2A and CN201920292985.8U) are used to deliver cooling airflow to different regions such as the leading edge, middle chord, and trailing edge. At the leading edge of the blade, a combination of impingement cooling and film cooling is commonly used (such as CN201811613132.6A) to cope with the highest heat load. In areas with high cooling difficulty, such as the trailing edge, turbulence column arrays (such as CN210118169U) or slotted structures are widely used to enhance heat transfer and gas exhaust. In addition, significant progress has been made in the refined design of cooling structures, including optimizing the azimuth and tilt angles of cooling holes (such as CN201920292985.8U) and designing non-tail-edge expanding cooling slots (such as CN109653806A) to improve air film adhesion and reduce the mixing loss of cooling airflow with the mainstream.

[0027] Although gas-thermal coupling numerical simulations (such as CN202410556042.7A) can optimize fixed structures, they cannot fundamentally solve the problems of dynamic thermal loads and adaptability to varying operating conditions faced by blades in real working environments. Therefore, existing turbine guide vane cooling technologies are facing a common technical bottleneck in their pursuit of higher performance: how to enable the cooling system to overcome the limitations of static design and achieve adaptive intelligent matching and response to thermal loads across the entire operating range of the engine.

[0028] Based on the problems existing in the prior art, the present invention proposes an adaptive cooling structure for aero-engine turbine guide vanes. The aero-engine turbine guide vane includes a blade body, and at least one cooling cavity is provided inside the blade body. An adaptive adjustment unit is provided on the cooling cavity. The adaptive adjustment unit dynamically adjusts the flow rate or direction of the cooling medium flowing through the cooling cavity based on the thermal load changes under the operating conditions of the turbine guide vane, so as to achieve adaptive control of the cooling intensity of the blade body.

[0029] Preferably, the adaptive adjustment unit includes a flow regulating component disposed circumferentially in at least one cooling chamber, which is capable of changing its own opening or flow area in response to temperature or pressure changes in the area surrounding the cooling chamber. Preferably, the cooling cavity has multiple impact slits extending through the blade span along the circumferential direction, and the lip of the impact slit is formed by a flow regulating component; the flow regulating component is a shape memory alloy actuator, and when the local temperature rises to the phase transition trigger temperature of the shape memory alloy, the shape memory alloy actuator deforms, thereby increasing the effective flow area of ​​the impact slit.

[0030] Preferably, the shape memory alloy actuator is a bimetallic strip, which is respectively formed or connected to the walls on both sides of the impact slit; the bimetallic strip undergoes bending deformation based on temperature changes, thereby adjusting the opening of the impact slit.

[0031] Preferably, the bimetallic strip is made of gold.

[0032] This invention also proposes a design method for an adaptive cooling structure for aero-engine turbine guide vanes, comprising the following steps: Step 1: Define the effective uniformly distributed load ; The effective uniformly distributed load To represent the ideal uniformly distributed load equivalent to the complex pressure distribution generated by the jet airflow from the impact slit, it is determined by the layout P, ​​number N, width W, and total pressure at the main inlet of the impact slit. and the total temperature of the main entrance The function is represented as:

[0033] Step 2: Calculate the deformation of the metal sheet in the direction perpendicular to it. y z ; According to the effective uniformly distributed load The deformation was calculated using the beam bending deformation formula. y z :

[0034] in, l denoted as , where E is the length of the metal sheet; E is the elastic modulus of the metal sheet; and I is the moment of inertia of the cross section. Step 3: Calculate the deformation of the metal sheet in the direction parallel to it. y x ; Considering the distributed torque T generated simultaneously by the gas impact, first calculate the torsion angle at the end of the support plate. :

[0035] in, G It is the shear modulus of the material. J It is the torsional moment of inertia of the cross section; Subsequently, based on the stated torsion angle and vertical deformation y z Calculate the deformation in the parallel direction. y x :

[0036] Step 4: Establish the relationship between deformation and changes in slit cross-sectional area to achieve adaptive flow rate adjustment; Based on the calculated deformation in the parallel direction Determine the change in the effective flow cross-sectional area of ​​the impact slit, wherein the change in cross-sectional area is related to... Positive correlation; by iteratively optimizing the layout P, ​​number N, and width W parameters of the impact slits, the cross-sectional area variation can adapt to the total pressure at the main inlet. and the total temperature of the main entrance The changes in temperature allow for self-regulation of the cooling flow rate.

[0037] This invention also proposes a parametric optimization design method for an adaptive cooling structure of aero-engine turbine guide vanes, the specific steps of which are as follows: Step 1: Establish analysis baseline Based on the three-dimensional model of the turbine guide vane including the initial impact slit layout, flow field simulation analysis is performed to obtain the baseline heat load distribution and cooling efficiency distribution on the blade surface. Step 2: Set optimization objectives and constraints The optimization objective is to improve overall cooling efficiency, and the core constraint is that the maximum surface temperature of the blades is lower than the allowable temperature of the material. Step 3: Execute iterative optimization loop Repeat the following sub-steps until the optimization objective and constraints set in step 2 are met: Step 3.1: Simulation and Analysis: Perform flow field simulation on the current version of the turbine guide vane model to obtain its temperature field and cooling efficiency data; Step 3.2: Strategy Formulation: Compare the current simulation results with the baseline heat load distribution to identify areas with insufficient cooling and areas with excessive cooling, and formulate adjustment strategies for the layout, number, and / or width of the impact slits based on this. Step 3.3: Model Update: Based on the adjustment strategy, the geometric parameters of the impact slit in the turbine guide vane model are modified parametrically to generate a new model for the next iteration.

[0038] The present invention also proposes an aero-engine in which the high-pressure turbine guide vane section adopts the adaptive cooling structure of the aero-engine turbine guide vane.

[0039] The above technical solution will be further explained below with reference to the accompanying drawings: In one embodiment, refer to Figure 1 As shown, this invention designs an adaptive cooling structure for turbine guide vanes. The adjustable cooling structure of the blade mainly focuses on the internal impact cooling structure: the impact slit. The plates on both sides of the impact slit are made of a material with a special elastic modulus, approximately within the range of 50GPa-70GPa, allowing it to automatically adjust the width of the slit according to changes in the guide vane inlet parameters. This influences the impact effect of the cooling air and adjusts the distribution ratio of the cooling air, solving the problem of a sharp decrease in the cooling efficiency of turbine guide vanes under high operating conditions. Gold is preferably used as the material for the plates on both sides of the impact slit.

[0040] The plates on both sides of the impact slit are defined as impact plates. The specific deformation of the impact plates follows the mechanical behavior of a cantilever beam, and its end deformation w is related to the load it receives and its own length. The elastic modulus E of the material is related to the moment of inertia I of the cross section.

[0041] Here, a core concept is defined as "effective uniformly distributed load" ( Although the actual pressure distribution generated by the slit jet is complex, the resulting end-effector deformation can be equivalent to the deformation produced by an ideal uniformly distributed load. This equivalent ideal load is called the "effective uniformly distributed load" (AEP). )". "The layout P, ​​number N, width W, and total pressure of the main inlet of the impact slits" and the main entrance total temperature Related. The specific function expression is:

[0042] Therefore, the deformation perpendicular to the support plate direction Calculated from equation (1): (1) In the formula, For an effective uniformly distributed load, i.e., gas pressure;l E is the length of the impact support plate; E is the elastic modulus (Young's modulus) of the support plate material; I is the moment of inertia of the cross section.

[0043] The gas impact not only generated forces that caused the beam to bend. This also generates a moment T that causes the beam to twist, resulting in an end twist angle along the direction parallel to the support plate: (2) Where T is the torque distributed per unit length, G is the shear modulus of the material, and J is the torsional moment of inertia of the cross section.

[0044] The deformation along the direction parallel to the support plate can then be approximately calculated using equation (3): (3) Therefore, the change in the cross-sectional area of ​​the impact slit is related to A positive correlation allows for adaptive flow adjustment.

[0045] In one embodiment, refer to Figure 5 As shown, a parametric optimization design method for an adaptive cooling structure of turbine guide vanes is presented. This method transforms the traditional experience-based design process into a closed-loop iterative optimization process based on fluid-structure-thermal coupling simulation, thereby obtaining the optimal impact slit layout, number, and slit width. Its core is a parametric optimization design method for slit layout based on target heat load. The specific steps are as follows: First, a turbine guide vane model incorporating the initial impact slit layout was established using 3D modeling software. Then, CFD simulation was employed to analyze this model and obtain the heat load distribution map of the blade surface, serving as the fundamental basis for all subsequent optimization adjustments. Based on this, a clear optimization objective function (improving overall cooling efficiency) was defined. (5%), and simultaneously meet the core constraint conditions that the blade surface temperature is below the allowable limit or the cooling air consumption is less than the preset upper limit.

[0046] For each subsequent design iteration, CFD simulation is first performed on the current scheme. The simulated temperature contour map is compared with the baseline heat load map to identify high-heat areas with insufficient cooling or over-cooling areas with wasted cool air. Key data points are also extracted, such as the specific temperature values ​​of the hottest spots at the leading edge. Based on this analysis, the next adjustment strategy is formulated. For example, increasing the width of the SH1 and SH2 slots can enhance the impact cooling of the leading edge stagnation area, or adjusting the relative positions of SS1, SS2, and SS3 can optimize the coverage of the high-heat load zone on the suction surface.

[0047] Subsequently, according to the established strategy, the layout, number, or width parameters of the impact slits are manually modified in UG software to generate a new design model, which is then subjected to CFD simulation again. This "analysis-decision-adjustment" cycle will be repeated, and key indicators (such as overall cooling efficiency and maximum wall temperature) in each iteration will be recorded until the indicators reach the preset optimization target. Therefore, the final accurate impact slit layout and dimensions are obtained.

[0048] In one embodiment, refer to Figure 1 , Figure 2 As shown, with the goal of improving the overall cooling efficiency of the deformed structure by 5% compared to the undeformed structure, through repeated iterations, the impact slot width that can ultimately achieve a 5% improvement in overall cooling efficiency was obtained. Within the two impact chambers of the blade, Figure 1 The front cavity shown in (a) has five impact slits: SH1=0.53mm, SH2=0.65mm, SH3=0.518984mm, SH4=0.518984mm, and SH5=0.659680mm. Figure 1 As shown in (b), the rear cavity has two impact slits on the pressure side and three impact slits on the suction side, with PS1=0.44267mm, PS2=0.45700mm; SS1=0.27829mm, SS2=0.30289mm, and SS3=0.29890mm. All impact slits are designed as through structures along the blade span.

[0049] The adaptive adjustment described in this invention is not achieved by adding moving parts, but by a clever combination of optimized fixed geometry and principles of fluid mechanics and thermodynamics.

[0050] Its working principle is as follows: Under high and low operating conditions, there is an absolute driving pressure difference between the inner and outer sides of the impact plate, which forces the plates on both sides of the impact slit to undergo slight, controllable deformation. The amount of deformation is proportional to the driving pressure difference. Under low operating conditions, because the inlet pressure itself is low, the corresponding absolute pressure added to the inner and outer sides of the impact plate is also low, so the deformation of the baffles on both sides of the impact slit is small; while under high operating conditions, the inlet pressure of the primary and secondary flows is at a higher level, and the corresponding absolute pressure added to the inner and outer sides of the impact plate is also higher, so the deformation of the baffles on both sides of the impact slit is larger, such as... Figure 2 (a) shows the mid-section structure of the impact slit before deformation. Figure 2 (b) is a schematic diagram of the mid-section structure after the impact slit is deformed. Correspondingly, under low operating conditions, the impact slit is still in a narrow state with large internal flow resistance and small cold air flow. Conversely, under high operating conditions, the impact slit is wider with smaller internal flow resistance and larger cold air flow, which effectively improves the overall cooling efficiency and ultimately maintains high-efficiency cooling across the entire operating range.

[0051] This invention uses computational fluid dynamics (CFD) simulation, selecting typical operating conditions of the engine during high-altitude cruise. Specific boundary conditions are set as follows: total pressure of the turbine guide vane mainstream is 2026.5 kPa, total temperature of the high-temperature combustion gas is 2000 K, and the inlet Mach number is 0.3. To simulate the cooling system inside the blades, the pressure ratio of the cooling air to the mainstream combustion gas is set to 1.1 to ensure effective injection of cooling airflow into the mainstream channel; simultaneously, the temperature ratio is set to 2.0, meaning the temperature of the cooling air is 1000 K.

[0052] Reference Figure 3 and Figure 4 The diagram shows the distribution cloud map of the overall cooling effect of the blade structure before and after deformation, as well as the distribution of the overall cooling effect at the mid-section. Observation of the distribution of the overall cooling effect at the mid-section of the guide vane confirms that adaptive cooling technology can improve the overall cooling effect of the blade. Calculations show that the ratio of primary to secondary flow rates before deformation was 2.567%, and the average overall cooling effect was 0.522457. After deformation, the ratio of primary to secondary flow rates was 2.4458%, and the average overall cooling effect was 0.551006. The overall cooling effect increased by 5.46% compared to the original structure.

[0053] Based on the technical solution of this invention, the fixation defects of existing geometric impact cooling structures are solved, as follows: 1. Address the core deficiency of being unable to adapt to changes in operating conditions. The primary technical problem this invention aims to solve is: how to enable the impact cooling structure of turbine blades to break free from the constraints of fixed geometry, allowing it to intelligently respond to changes in engine inlet parameters (such as pressure and temperature) and automatically adjust its cooling characteristics, thereby maintaining efficient and appropriate cooling intensity throughout the entire flight envelope from takeoff, climb, cruise to landing, and completely changing the passive situation where existing structures can only work at a single design point.

[0054] 2. Solve the problem of insufficient cooling capacity under high altitude and low operating conditions. This invention designs a structure that can automatically adjust to changes in environmental pressure, maintaining or optimizing the flow state of cooling gas even at high altitudes. This ensures that the impact heat transfer effect does not significantly distort or decrease during high-altitude cruising, guarantees effective coverage of the cooling film, and improves the reliability of the blades under low operating conditions.

[0055] 3. Solve the problem of high manufacturing and maintenance costs. Despite employing high-performance materials, this invention aims to address cost issues through innovative structural principles: how to replace or simplify the extremely complex, precise, and expensive multi-channel cooling systems and their manufacturing processes in existing technologies with a simple, reliable, and self-adaptive structure that requires no external control source. By reducing reliance on complex irregular holes and micro-hole arrays, manufacturing difficulty is reduced, and failures such as micro-hole blockage are avoided in principle, thereby potentially reducing manufacturing and maintenance costs throughout the entire lifecycle.

[0056] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. An adaptive cooling structure for an aero-engine turbine guide vane, the aero-engine turbine guide vane comprising a blade body, wherein at least one cooling chamber is disposed inside the blade body; characterized in that: An adaptive adjustment unit is provided on the cooling chamber. The adaptive adjustment unit dynamically adjusts the flow rate or direction of the cooling medium flowing through the cooling chamber based on the thermal load changes of the turbine guide vane under its operating conditions, so as to achieve adaptive control of the cooling intensity of the blade.

2. The adaptive cooling structure for aero-engine turbine guide vanes according to claim 1, characterized in that: The adaptive adjustment unit includes a flow regulation component disposed circumferentially in at least one cooling chamber. The flow regulation component is capable of changing its opening or flow area in response to temperature or pressure changes in the area surrounding the cooling chamber.

3. The adaptive cooling structure for aero-engine turbine guide vanes according to claim 2, characterized in that: The cooling chamber has multiple impact slits extending through the span of the blades along its circumferential direction. The lips of the impact slits are formed by flow regulating components. The flow regulating components are shape memory alloy actuators. When the local temperature rises to the phase transition trigger temperature of the shape memory alloy, the shape memory alloy actuator deforms, thereby increasing the effective flow area of ​​the impact slits.

4. The adaptive cooling structure for aero-engine turbine guide vanes according to claim 3, characterized in that: The shape memory alloy actuator is a bimetallic strip, which is respectively formed or connected to the walls on both sides of the impact slit; the bimetallic strip undergoes bending deformation based on temperature changes, thereby adjusting the opening of the impact slit.

5. The adaptive cooling structure for aero-engine turbine guide vanes according to claim 4, characterized in that: The bimetallic strip is made of gold.

6. The adaptive cooling structure for aero-engine turbine guide vanes according to claim 5, characterized in that: The cooling chamber inside the blade includes a front chamber and a rear chamber; The wall of the front cavity is provided with five impact slits along the circumferential direction, and the widths of the bimetallic strips are: SH1=0.53mm, SH2=0.65mm, SH3=0.518984mm, SH4=0.518984mm, SH5=0.659680mm respectively. The rear cavity has five impact slits along its circumferential direction on its wall surface. Two slits are provided on the pressure side, with bimetallic strip widths of PS1=0.44267mm and PS2=0.45700mm respectively; and three slits are provided on the suction side, with bimetallic strip widths of SS1=0.27829mm, SS2=0.30289mm and SS3=0.29890mm respectively.

7. A design method for an adaptive cooling structure for aero-engine turbine guide vanes according to any one of claims 1-6, characterized in that, Includes the following steps: Step 1: Define the effective uniformly distributed load ; The effective uniformly distributed load To represent the ideal uniformly distributed load equivalent to the complex pressure distribution generated by the jet airflow from the impact slit, it is determined by the layout P, ​​number N, width W, and total pressure at the main inlet of the impact slit. and the main entrance total temperature The function is represented as: Step 2: Calculate the deformation of the metal sheet in the direction perpendicular to it. y z ; According to the effective uniformly distributed load The deformation was calculated using the beam bending deformation formula. y z : in, l denoted as , where E is the length of the metal sheet; E is the elastic modulus of the metal sheet; and I is the moment of inertia of the cross section. Step 3: Calculate the deformation of the metal sheet in the direction parallel to it. y x ; Considering the distributed torque T generated simultaneously by the gas impact, first calculate the torsion angle at the end of the support plate. : in, G It is the shear modulus of the material. J It is the torsional moment of inertia of the cross section; Subsequently, based on the stated torsion angle and vertical deformation y z Calculate the deformation in the parallel direction. y x : Step 4: Establish the relationship between deformation and changes in slit cross-sectional area to achieve adaptive flow rate adjustment; Based on the calculated deformation in the parallel direction Determine the change in the effective flow cross-sectional area of ​​the impact slit, wherein the change in cross-sectional area is related to... Positive correlation; by iteratively optimizing the layout P, ​​number N, and width W parameters of the impact slits, the cross-sectional area variation can adapt to the total pressure at the main inlet. and the main entrance total temperature The changes in temperature allow for self-regulation of the cooling flow rate.

8. A parametric optimization design method for an adaptive cooling structure of an aero-engine turbine guide vane as described in any one of claims 1-6, characterized in that... The specific steps are as follows: Step 1: Establish analysis baseline Based on the three-dimensional model of the turbine guide vane including the initial impact slit layout, flow field simulation analysis is performed to obtain the baseline heat load distribution and cooling efficiency distribution on the blade surface. Step 2: Set optimization objectives and constraints The optimization objective is to improve overall cooling efficiency, and the core constraint is that the maximum surface temperature of the blades is lower than the allowable temperature of the material. Step 3: Execute iterative optimization loop Repeat the following sub-steps until the optimization objective and constraints set in step 2 are met: Step 3.1: Simulation and Analysis: Perform flow field simulation on the current version of the turbine guide vane model to obtain its temperature field and cooling efficiency data; Step 3.2: Strategy Formulation: Compare the current simulation results with the baseline heat load distribution to identify areas with insufficient cooling and areas with excessive cooling, and formulate adjustment strategies for the layout, number, and / or width of the impact slits based on this. Step 3.3: Model Update: Based on the adjustment strategy, the geometric parameters of the impact slit in the turbine guide vane model are modified parametrically to generate a new model for the next iteration.

9. An aircraft engine, characterized in that, Its high-pressure turbine guide vane section adopts the adaptive cooling structure for aero-engine turbine guide vanes as described in any one of claims 1-6.

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