A multi-stage axial flow compressor considering inner cooling of last stage and design method thereof

Abstract

The application discloses a kind of multi-stage axial flow compressor considering end stage internal cooling and design method.First, the distribution of internal pressure field and temperature field of multi-stage axial flow compressor is obtained by numerical calculation, combined with the thermodynamic properties of blade material, determine the number and position of cooling stage.Second, cooling air flow is introduced from a certain position of the front stage, sequentially enters hub and rotor / stator internal cooling channel, realizes blade cooling.After that, cooling gas returns to upstream, forms closed cooling circuit.To avoid cracks on the surface of blade and cause local high temperature air flow to pour into cooling channel, a cooling technology with "blade embedded heat pipe + loop heat exchanger" as the core in limited space is proposed, heat pipe transfers heat inside the blade to loop heat exchanger in hub / casing, and heat is carried away by cooling air flow.The above scheme not only strengthens the cooling effect of blade, but also guarantees the structural reliability, reduces flow loss, realizes the double optimization of end stage blade in aerodynamic and structure.

Description

Technical Field

[0001] The present invention belongs to the technical field of aero-engine compressor design and manufacturing, and relates to aerodynamic and structural performance optimization technology for the last-stage blades of aero-engine compressor components. Specifically, it relates to a multi-stage axial flow compressor and a design method that take into account the internal cooling of the last stage. The method is used to solve the problems of high heat load and high aerodynamic loss of the last-stage blades of a high-pressure ratio multi-stage axial flow compressor, thereby achieving a dual improvement in the aerodynamic performance and structural reliability of the compressor. Background Art

[0002] Modern aircraft engines are constantly developing towards high thrust, low fuel consumption, and high reliability, placing increasingly stringent demands on the aerodynamic and structural design of compression components. For aircraft engines, higher pressure ratios mean higher efficiency and lower fuel consumption. Therefore, current high-thrust-to-weight ratio aircraft engines are typically equipped with multi-stage axial-flow compression systems, achieving pressure ratios exceeding 50. Within a multi-stage axial-flow compression system, the air temperature rises by approximately 90°C with each stage of compression. As the number of stages and pressure ratios in multi-stage axial-flow compressors increases, the air temperature around the last-stage blades rises rapidly. Excessive heat loads can cause blade damage. Furthermore, the small size of the last-stage blades (5-10 mm) leads to significant secondary flow, which dramatically increases aerodynamic losses. Therefore, for the last-stage blades of high-pressure multi-stage axial-flow compressors, two technical bottlenecks must be overcome: structural failure induced by excessive heat loads, and the regulation of secondary flow in the end zone due to size effects.

[0003] Traditional compressor designs usually do not consider the wall heat transfer effect, and the blade / hub surface is assumed to be insulated. This design concept can still meet the needs of medium and low pressure ratio compressors. However, for the last-stage blades of high-pressure ratio multi-stage axial flow compressors under severe gas-thermal conditions, it is urgent to break through the adiabatic wall assumption and introduce the cooling effect into the design system. As an effective thermal flow / aerodynamic control method, wall cooling has great potential in improving the structure and aerodynamic performance of the last-stage blades of multi-stage compressors. On the one hand, wall cooling can effectively reduce the temperature of the blade surface, thereby reducing the blade thermal load and extending the service life; on the other hand, wall cooling has a significant impact on the development of the blade / endwall boundary layer, thereby regulating the end-stage load and low-energy vortex force characteristics of the last-stage blade, weakening the unsteady interference intensity between the blade boundary layer and the secondary flow in the end area, and reducing aerodynamic losses.

[0004] At present, research on compressor blade cooling mainly focuses on internal cooling and external cooling technologies. Internal cooling technology reduces blade temperature by designing a complex cooling channel network inside the blade and utilizing the convective heat exchange between cooling air and the internal surface of the blade. External cooling technology forms a layer of cooling air film by setting air film holes or spray cooling on the external surface of the blade to isolate the high-temperature airflow from direct contact with the blade surface. For example, Chinese invention patent application CN115898957A discloses a compressor rotor cooling structure adapted to high Mach number inflow conditions. It reduces the surface temperature by setting cooling air grooves on the blade surface and introducing cooling airflow. However, this technology only addresses the cooling problem of the blade surface, and does not solve the problem of designing complex cooling channels inside the blade, nor does it optimize the unsteady interaction between the blade boundary layer and the secondary flow in the end area. For example, Chinese invention patent CN116517885B discloses a cooling structure for the rear stage casing of a compressor, which uses cooling pipes and cooling wires wrapped around them for cooling to improve compressor efficiency. However, this technology only addresses the cooling needs of the rear stage casing of the compressor and fails to solve the structural failure and aerodynamic loss problems of the last stage blades under high thermal loads.

[0005] In summary, the last-stage blades of high-pressure-ratio multi-stage axial-flow compressors face multiple challenges, including a harsh thermal environment, small size, and high losses induced by intense secondary flow. To ensure stable and efficient operation of multi-stage compressors, a design method for last-stage blades that considers wall cooling is urgently needed. Achieving efficient blade cooling within confined spaces while ensuring structural reliability and significant flow field loss reduction is a key issue that needs to be addressed. Summary of the Invention

[0006] (1) Purpose of the invention

[0007] In view of the defects of the existing high-pressure ratio multi-stage axial flow compressor, such as the excessive heat load of the last-stage blades leading to structural failure and the strong secondary flow in the end area inducing significant aerodynamic losses, in order to solve at least one of the above-mentioned and other technical problems in the prior art, the purpose of the present invention is to provide a multi-stage axial flow compressor and a design method that take into account the cooling of the last-stage blades, break through the inherent assumption of the adiabatic wall conditions in the traditional design method, and propose a cooling scheme with "blade embedded heat pipe + loop heat exchanger" as the core in a confined space by introducing the wall cooling effect. This scheme enhances the cooling effect, effectively reduces the heat load of the last-stage blades, ensures structural strength, and at the same time weakens the interference intensity between the blade surface boundary layer and the secondary flow, reduces the turbulent pulsation generation rate, and realizes dual regulation of the heat load and aerodynamic losses of the last-stage blades, thereby solving the bottleneck problem of high heat load and high loss of the last-stage blades of the multi-stage axial flow compressor.

[0008] (2) Technical solution

[0009] In order to achieve the purpose of the invention and solve the technical problems, the present invention adopts the following technical solutions:

[0010] The first object of the present invention is to provide a design method for a multi-stage axial flow compressor that takes into account internal cooling of the last stage. The design method is used for cooling the last stage blades, thereby reducing blade thermal load and aerodynamic losses, improving the structural strength and aerodynamic performance of the blades, and ensuring efficient and stable operation of the compressor. The implementation of the design method includes at least the following steps:

[0011] SS1. Structural-Aerodynamic Coupling Analysis:

[0012] A structural and aerodynamic coupling analysis model was established. By adjusting the speed and flow coefficient of the multi-stage compressor, the internal pressure and temperature field distributions of the compressor under different operating conditions were obtained. Combined with the thermodynamic properties of the blade material, the number of blade stages that required cooling and the specific cooling area were determined.

[0013] SS2. Design of a closed cooling circuit for rotor and stator blades:

[0014] A cooling airflow is drawn out from a certain starting position of the front stage, passes through the first cooling channel inside the hub, the cooling channel inside the rotor blade, and then returns to a certain upstream recovery position through the second cooling channel inside the hub, thereby constructing a rotor blade cooling circuit; a cooling airflow is drawn out from a certain starting position of the front stage, passes through the first cooling channel inside the casing, the cooling channel inside the stator blade, and then returns to a certain upstream recovery position through the second cooling channel inside the casing, thereby constructing a stator blade cooling circuit;

[0015] SS3. Optimize the cooling circuit starting and recovery locations:

[0016] Based on the pressure field distribution inside the compressor, a location with sufficient pressure to support the cooling airflow to complete the entire loop is selected as the initial starting position, and a location where the recovered airflow pressure exceeds the local mainstream pressure is selected as the initial recovery position. Through fluid-structure interaction calculations, the velocity and temperature field information at various locations inside the compressor under the cooling loop scheme is updated, and the starting and recovery positions are iteratively adjusted until the preset blade cooling amplitude and the cooling airflow are smoothly recovered without backflow are met.

[0017] SS4. Design of internal cooling structure for blades in confined spaces:

[0018] A composite cooling structure including heat pipes and a loop heat exchanger is provided in the rotor / stator blades, casing, and hub that require cooling, wherein the heat pipes are respectively embedded in the rotor and stator blades, and the evaporation ends of the heat pipes are arranged in the high heat load area of ​​the blades. The loop heat exchangers are respectively provided in the hub and casing, and the hot side of the loop heat exchanger is tightly connected to the condensation end of the heat pipe, and the cold side of the loop heat exchanger is respectively connected to the first cooling channel inside the hub and the first cooling channel inside the casing;

[0019] SS5. Iterative Optimization and Verification of Cooling Solutions:

[0020] The cooling scheme is iteratively optimized through numerical simulation and / or experimental testing to evaluate the impact of the cooling design on the blade surface temperature distribution, boundary layer development, and end-zone aerodynamic losses. The relevant parameters of the internal cooling channels, heat pipes, and / or loop heat exchangers of the hub and casing are optimized and adjusted to verify whether the cooling design meets the requirements of the blade cooling range and smooth circulation of the cooling airflow, thereby forming a final cooling system design scheme.

[0021] The second object of the present invention is to provide a multi-stage axial flow compressor that takes into account the internal cooling of the last stage. The compressor is designed and optimized using the above-mentioned multi-stage axial flow compressor design method that takes into account the internal cooling of the last stage, thereby achieving efficient cooling of the last-stage blades and improving aerodynamic performance.

[0022] (3) Technical effects

[0023] Compared with the prior art, the multi-stage axial flow compressor and design method provided by the present invention, which takes into account the internal cooling of the last stage, overcome the technical bottlenecks of high heat load-induced structural failure of the last stage blades and efficient control of the secondary flow in the lower end area due to size effect. Specifically, it has the following beneficial and significant technical effects:

[0024] (1) In view of the small size and compact structure of the last-stage blades of a multi-stage axial compressor, the present invention proposes a cooling technology based on "blade-embedded heat pipe + loop heat exchanger" in a confined space. The above scheme utilizes the efficient thermal conductivity of the heat pipe to quickly transfer the heat from the high-heat-load area inside the blade to the loop heat exchanger in the hub or casing, and then removes the heat through the cooling airflow. Compared with the traditional blade internal cooling channel scheme, on the one hand, it simplifies the cooling channel structure, avoids the problem of blade surface cracks induced by complex internal cooling channel design, and improves the overall structural reliability of the blade; on the other hand, through the close coupling of the loop heat exchanger and the heat pipe, efficient heat transfer is achieved, the cooling effect is enhanced, the heat load of the last-stage blade is effectively reduced, and the service life of the blade is extended.

[0025] (2) This invention breaks through the inherent assumption of adiabatic wall conditions in previous multi-stage compressor designs and introduces the wall cooling effect into the design system. By introducing the design concept of wall cooling, while avoiding excessive heat load on the blades that may lead to structural failure, it also weakens the interference intensity between the blade surface boundary layer and the secondary flow, thereby delaying the generation of turbulent pulsation in the end area and reducing the aerodynamic losses induced by the mixing of the blade boundary layer and the secondary flow in the end area, thus achieving loss reduction and efficiency improvement.

[0026] (3) The multi-stage axial flow compressor design method considering the final stage internal cooling proposed in this invention can evaluate the effect of wall cooling on the blade surface temperature distribution, boundary layer growth rate, and three-dimensional turbulent mixing in the end zone, and optimize and adjust the cooling structure parameters based on the evaluation results to achieve a dual improvement in the structural strength and aerodynamic performance of the final stage blades, thereby ensuring stable and efficient operation of the compressor. In addition, the present invention has good universality and scalability and can be applied to the final stage blade cooling design of multi-stage axial flow compressors with different loads and sizes, providing an effective technical solution for solving the cooling problem of compressor blades in high temperature and high pressure environments. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 This is a flow chart of the implementation of the multi-stage axial flow compressor design method considering the final stage internal cooling of the present invention;

[0028] Figure 2 Schematic diagram of the overall scheme of the multi-stage axial flow compressor cooling circuit in an embodiment of the present invention;

[0029] Figure 3 This is a schematic diagram of the cooling channel design of the "blade embedded heat pipe + loop heat exchanger" in the present invention;

[0030] Figure 4 Schematic diagram of the temperature distribution of the blade surface with and without cooling, where (a) is the temperature distribution under adiabatic conditions without cooling; (b) is the temperature distribution after adopting the internal cooling scheme. z / C ax represents the dimensionless chordal position of the blade, the ordinate x / H Indicates the dimensionless spanwise height of the blade, and the color of the cloud indicates the dimensionless temperature T / T 01 ;

[0031] Figure 5 Schematic diagram of velocity fluctuation and boundary layer displacement thickness of the blade section with and without cooling, where (a) is the result under adiabatic and no cooling conditions; (b) is the result after adopting the internal cooling scheme. z / C ax represents the dimensionless chordal position of the blade, the ordinate y / C ax Represents the dimensionless normal distance of the blade surface, and the color of the cloud represents the velocity pulsation intensity u ' RMS / U 1, the yellow dashed line represents the boundary layer displacement thickness distribution;

[0032] Figure 6With or without cooling near the lower end wall (leaf span position x / H =0.15) Reynolds normal stress comparison diagram, where (a) is the flow direction Reynolds normal stress distribution; (b) is the circumferential direction Reynolds normal stress distribution; (c) is the span direction Reynolds normal stress distribution. z / C ax represents the dimensionless chordal position of the blade, and the ordinates are the dimensionless flow direction ( ), circumferential ( ) and spanwise ( )Reynolds normal stress. DETAILED DESCRIPTION

[0033] The present invention aims to provide a multi-stage axial flow compressor design method that takes into account the cooling of the last-stage blades, thereby achieving dual control of the thermal load and aerodynamic losses of the last-stage blades. To further clarify the objectives, technical solutions, and advantages of the present invention, the technical solutions in the embodiments of the present invention will be described in more detail below in conjunction with the accompanying drawings. The described embodiments are only some, not all, of the embodiments of the present invention, and are exemplary and intended to explain the present invention. They should not be construed as limiting the present invention.

[0034] Example 1: Design method

[0035] The key to the multi-stage axial flow compressor design method of the present invention, which takes into account the cooling of the last-stage blades, lies in the cooling technology of the last-stage blades of the multi-stage axial flow compressor in a confined space. Figure 1 As shown, the main steps and corresponding solutions of the design method of the present invention when implemented are as follows:

[0036] SS1. Structural-Aerodynamic Coupling Analysis:

[0037] For a specific multi-stage axial flow compressor, a coupled structural and aerodynamic analysis model was first established, and numerical calculations were performed. By adjusting the speed and flow coefficient of the multi-stage compressor, the time-averaged pressure and temperature distributions within the compressor were determined under different operating conditions. Based on this pressure and temperature information, combined with the thermodynamic properties of the blade material, the number of blade stages requiring cooling and their specific locations were determined.

[0038] As a preferred method, the structural-aerodynamic coupling analysis model is based on three-dimensional modeling of the flow path and blade geometry of the multi-stage axial flow compressor. The RANS equations and turbulence model are used to solve the pressure field and temperature field distribution inside the compressor, analyze the temperature rise on the blade surface and end area, and the turbulent pulsation generation rate. At the same time, combined with the thermodynamic properties of the blade material, the number of stages and specific cooling areas that require cooling design are clearly determined.

[0039] SS2. Design of a closed cooling circuit for rotor and stator blades:

[0040] A cooling air flow is drawn out from a certain starting position of the front stage, passes through the first cooling channel inside the hub and the cooling channel inside the rotor blade in sequence, and then returns to a certain upstream recovery position through the second cooling channel inside the hub to construct a rotor blade cooling circuit; a cooling air flow is drawn out from a certain starting position of the front stage, passes through the first cooling channel inside the casing and the cooling channel inside the stator blade in sequence, and then returns to a certain upstream recovery position through the second cooling channel inside the casing to construct a stator blade cooling circuit.

[0041] Specifically, in order to minimize the mixing of the cooling airflow with the local mainstream, the present invention first proposes a closed cooling circuit for the stator blades, the overall scheme of which is as follows: Figure 2 As shown. For the closed cooling circuit of the rotor blade (R5), the cooling airflow is drawn out from a certain upstream position (position A), first enters the internal cooling channel of the hub ("channel 1"), and reaches the inlet of the rotor blade cooling channel (position B). The heat is removed through convection and impact processes inside the rotor (R5) blade channel, reducing the surface temperature of the rotor blade. After that, the cooling airflow passes through the internal cooling channel of the hub ("channel 2") again and returns to a certain upstream position (position C), thus forming a cooling circuit. For the closed cooling circuit of the stator blade (S5), there is a similar cooling process: the cooling airflow is drawn out from a certain upstream position of the casing (position D), passes through the internal cooling channel of the casing ("channel 3") and the stator blade in sequence, and then enters the internal cooling channel of the casing ("channel 4") again, and the cooling airflow recovery is completed at the upstream F of the casing.

[0042] Preferably, in the rotor blade cooling circuit, the outlet position and recovery position of the cooling airflow are both located on the hub of the front stage of the compressor, and the outlet position is located at the high-pressure downstream stage, and the recovery position is located at the low-pressure upstream stage; in the stator blade cooling circuit, the outlet position and recovery position of the cooling airflow are both located on the casing of the front stage of the compressor, and the outlet position is located at the high-pressure downstream stage, and the recovery position is located at the low-pressure upstream stage.

[0043] SS3. Optimize the cooling circuit starting and recovery locations:

[0044] After completing the overall design, the next step is to determine the specific locations where the cooling airflow starts and is recovered. For the starting position (A or D), the pressure there must be sufficient to support the smooth recovery of the cooling airflow after passing through the loop. For the recovery position (C or F), the pressure of the recovered airflow must be greater than the local mainstream pressure, otherwise it will cause backflow. Based on the above principles, combined with the pressure and temperature field distribution within the compressor, the starting and recovery positions of the cooling channel are first preliminarily set. Specifically, a position where the pressure is sufficient to support the cooling airflow to complete the entire loop flow is selected as the initial starting position, and a position where the recovered airflow pressure is greater than the local mainstream pressure is selected as the initial recovery position. Fluid-structure coupling calculations are then performed to update the velocity and temperature field information at various locations within the compressor for this cooling loop scheme, paying particular attention to whether there is backflow near the recovery position. Through forward design, the starting and recovery positions are iteratively adjusted until the basic requirements of the preset blade cooling amplitude and the smooth recovery of the cooling airflow without backflow are met.

[0045] As a preference, during the optimization process of the starting and recovery positions of the cooling circuit, the cooling circuit layout is optimized using fluid-solid coupling iterative calculations. With the goal of minimizing cooling air flow and minimizing losses along the way, the path length, cross-sectional shape, and geometric dimensions of the cooling channel are optimized and adjusted, while meeting the requirements of the preset cooling amplitude of the blades and the smooth recovery of the cooling airflow without backflow.

[0046] SS4. Design of internal cooling structure for blades in confined spaces:

[0047] A composite cooling structure including heat pipes and loop heat exchangers is set in the rotor / stator blades, casing and hub that need to be cooled, wherein the heat pipes are respectively embedded in the rotor and stator blades, and the evaporation ends of the heat pipes are arranged in the high heat load area of ​​the blades. The loop heat exchangers are respectively arranged in the hub and casing, and the hot side of the loop heat exchanger is tightly connected to the condensation end of the heat pipe, and the cold side of the loop heat exchanger is respectively connected to the first cooling channel inside the hub and the first cooling channel inside the casing.

[0048] Theoretically, the internal cooling channels of the last-stage blades of a multi-stage axial compressor can be of any form. However, different forms of cooling channels are directly related to the heat exchange effect, which in turn affects the blade's thermal load, structural strength, and related aerodynamic performance parameters. For rotor blades, the interior is a closed cooling channel, and the cooling airflow cannot diffuse and overflow from the blade surface. Therefore, the airflow pressure in the cooling channel inside the blade is generally lower than the airflow pressure near the blade surface. Precisely because of this, when cracks appear on the blade surface due to long service life or external impact, the high-temperature, high-pressure airflow on the blade surface may quickly flow into the internal cooling channel, seriously affecting the cooling effect. When complex cooling channels are opened inside the very small last-stage blades, it is more likely to induce structural failure phenomena such as cracking on the blade surface, which in turn brings a lot of uncertainty to the cooling effect.

[0049] In order to solve the difficulty in designing the cooling channel of the last stage blade in a confined space, the present invention proposes a design scheme of built-in heat pipes in the blades (taking the rotor as an example), such as Figure 3 As shown, a heat pipe (P) is embedded within the rotor blade (R5), transferring heat to a loop heat exchanger (H). When the cooling airflow from upstream channel 1 passes through the loop heat exchanger (H), it removes the heat exchanged by the heat pipe P. A similar cooling strategy can be applied to the stator blade S5. This "blade-embedded heat pipe + loop heat exchanger" approach not only simplifies the cooling path but also, through optimized heat pipe structure, further enhances the cooling effect, minimizing problems such as blade surface cracking and even structural failure.

[0050] Preferably, the heat pipes embedded in the blades can adopt a multi-point distributed arrangement scheme. According to the distribution of high heat load areas inside the last-stage blades, the evaporation ends of the heat pipes are arranged at multiple points inside the blades, and uniform heat transfer is achieved through joint heat exchange between the heat pipes. At the same time, the condensation end of the heat pipe is tightly connected to the hot side of the loop heat exchanger using high thermal conductivity materials, and the layout of the condensation end is adapted to the geometric shape of the loop heat exchanger to improve the efficiency of heat transfer from the heat pipe to the cooling airflow.

[0051] In addition, the material selection of the heat pipe must meet the following requirements at the same time: it has good thermal conductivity within the operating temperature range of the blade, a thermal expansion coefficient similar to that of the blade material, sufficient high-temperature strength and fatigue resistance, and the connection between the heat pipe and the blade adopts a high-temperature brazing process to ensure good thermal conductivity and structural reliability; the overall size of the loop heat exchanger must meet the space constraints of the hub or casing.

[0052] SS5. Iterative Optimization and Verification of Cooling Solutions:

[0053] The cooling scheme is iteratively optimized through numerical simulation and / or experimental testing to evaluate the impact of the cooling design on the blade surface temperature distribution, boundary layer growth rate and end zone turbulence pulsation. The relevant parameters of the internal cooling channels, heat pipes and / or loop heat exchangers of the hub and casing are optimized and adjusted to verify whether the cooling design meets the requirements of the blade cooling range and smooth circulation of the cooling airflow, thus forming the final cooling system design scheme.

[0054] Preferably, in step SS5, the numerical simulation adopts the unsteady RANS method to establish a full three-dimensional calculation model including blades, heat pipes and heat exchangers, and accurately predicts the heat transfer characteristics of the blade surface based on the SST turbulence model coupled with a high-precision transition format. The calculation results need to be iterated through fluid-heat coupling to obtain a converged solution of the blade surface temperature distribution and flow field characteristics.

[0055] In addition, when evaluating the impact of cooling design on blade surface temperature distribution, boundary layer development and aerodynamic losses in the end area, it mainly includes: obtaining the blade surface temperature distribution with and without cooling conditions through numerical simulation, focusing on the local temperature drop near the high heat load area of ​​the blade, and evaluating the cooling effect and uniformity of the cooling design; comparing the turbulent pulsation characteristics and boundary layer development laws of the suction surface of the blade mid-section with and without cooling conditions, and evaluating the role of the cooling effect in weakening the trailing edge separation of the blade mid-section and delaying the growth of the boundary layer; comparing the three-dimensional Reynolds normal stress at different axial positions near the end wall with and without cooling conditions, and evaluating the effect of the cooling effect in weakening the lateral / radial migration of the secondary flow in the end area, reducing the three-dimensional mixing of turbulent flow and the total pressure loss of the outlet section.

[0056] When iteratively optimizing the cooling scheme based on the evaluation and analysis results, the geometric parameters of the hub and casing cooling channels, the layout of the heat pipes, and the design of the loop heat exchanger are adjusted to ensure a synergistic improvement in the cooling effect and the aerodynamic performance of the blades. Specifically, this includes: optimizing the path length, cross-sectional shape, and geometric dimensions of the cooling channels to achieve minimum cooling air flow and minimum along-the-path loss and improve flow uniformity; optimizing the material, diameter, length, and layout of the heat pipes to improve heat transfer efficiency and ensure structural stability under high-speed rotation conditions; optimizing the heat transfer area, flow channel structure, and installation position of the loop heat exchanger to achieve maximum heat transfer and reduce total pressure loss within a limited space; and the final cooling scheme should meet the requirements of blade cooling amplitude, boundary layer growth rate control, average total pressure loss coefficient control of the outlet section, and smooth circulation of the cooling airflow.

[0057] Example 2: Effect Verification

[0058] On the basis of the above-mentioned Example 1, as an example to further verify the effect of the technical solution of the present invention, Example 2 takes a certain cooling scheme as an example, evaluates the cooling effect of the last-stage blade through numerical simulation, and specifically analyzes the influence of the cooling scheme on the blade surface temperature distribution, boundary layer growth rate and aerodynamic performance. The results show that the internal cooling design of the last-stage blade can not only effectively reduce the heat load and extend the service life, but also is expected to improve the local flow characteristics.

[0059] The leaf surface temperature distribution with and without internal cooling is as follows Figure 4 As shown in (a) and (b), it can be seen that without cooling design, the blade surface temperature is z / C ax =0.5 and leaf position x / H =0.35 is higher ( T / T 01>0.76), the heat flux density in the local area is relatively large. After adopting the internal cooling solution, the leaf surface temperature in this area is significantly reduced under the cooling effect ( T / T 01 ≈0.64), axial position z / C ax =0.5, leaf position x / H =0.35, the local temperature drops by 14.2%, and the temperature distribution on the entire blade surface becomes more uniform, indicating that the internal cooling scheme can effectively reduce the heat load of the blade, especially the high-temperature area has a significant local cooling effect.

[0060] In order to reveal the effect and regulation mechanism of blade internal cooling on aerodynamic performance, Figure 5 The velocity fluctuation of the suction surface of the stator blade with and without cooling conditions was compared ( u ' RMS / U 1) and boundary layer displacement thickness ( δ TE *). It can be seen that under the adiabatic wall condition, open flow separation occurs at the trailing edge of the blade suction surface, and strong turbulent mixing induces high-level velocity fluctuations. z / C ax >0.6, a large area of ​​high horizontal velocity pulsation appeared in the downstream area ( u ' RMS / U 1>0.3), the boundary layer displacement thickness increases sharply, and the boundary layer displacement thickness at the trailing edge of the blade δ TE *reach 0.08 C ax When the cooling effect exists, although open separation still exists at the trailing edge of the blade, the separation scale is reduced, the high-level turbulence pulsation area is reduced, the growth of the boundary layer in the turbulent zone is also delayed, and the boundary layer displacement thickness at the trailing edge is reduced to 0.067 C ax Compared with the adiabatic condition, the internal cooling scheme leads to a 16.25% reduction in the boundary layer displacement thickness near the trailing edge.

[0061] In addition to regulating the growth of the boundary layer in the blade mid-section and the rate of turbulent pulsation generation, wall cooling also affects the spatiotemporal migration of the secondary flow near the end wall and the intensity of its unsteady interference with the local blade boundary layer, thereby changing the loss level near the end wall. Figure 6The Reynolds normal stress distribution at a section 15% of the blade height from the end wall is compared with and without cooling conditions. In the figure, (a) is the flow direction Reynolds normal stress distribution, (b) is the circumferential direction Reynolds normal stress distribution, and (c) is the span direction Reynolds normal stress distribution. 50% axial chord length ( z / C ax =0.5), the Reynolds normal stresses in all three directions are at extremely low levels. Downstream of 50% of the axial chord length, the transversely migrating secondary flow begins to roll up spanwise and strongly interferes with the blade suction surface boundary layer at that blade height, inducing strong turbulent pulsations. The Reynolds normal stresses in all three directions begin to increase dramatically, and losses increase. Compared to adiabatic conditions, the cooling effect not only reduces the growth rate of the blade surface boundary layer but also weakens the lateral / radial migration of the secondary flow near the end region by regulating the transition process and load distribution. As a result, the intensity of the interference between the blade boundary layer and the secondary flow near the end region is weakened, the Reynolds normal stresses in all three directions downstream of 50% of the axial chord length decrease significantly, and the average total pressure loss coefficient of the outlet section decreases by 13.3%. These results further demonstrate that internal blade cooling can not only reduce the thermal load of the last stage of a multi-stage compressor and extend its service life, but also significantly improve the three-dimensional flow performance near the end region, achieving loss reduction and efficiency improvement.

[0062] The above embodiments fully and effectively achieve the objectives of the present invention. Those skilled in the art will appreciate that the present invention includes, but is not limited to, the contents described in the accompanying drawings and the above specific embodiments. Although the present invention has been described with reference to the embodiments currently considered to be the most practical and preferred, it should be understood that the present invention is not limited to the disclosed embodiments, and any modifications that do not deviate from the functional and structural principles of the present invention are intended to be included within the scope of the claims.

Claims

1. A design method for a multi-stage axial flow compressor taking into account the final stage internal cooling, characterized in that: The design method comprises at least the following steps when implemented: SS1. Build a coupled structural and aerodynamic analysis model. By adjusting the speed and flow coefficient of a multi-stage compressor, determine the internal pressure and temperature distributions under different operating conditions. Combined with the thermodynamic properties of the blade material, determine the number of blade stages requiring cooling and the specific cooling areas. SS2. Cooling air is drawn from a starting point in the preceding stage, passes sequentially through the first cooling channel in the hub, the cooling channel in the rotor blades, and then returns to an upstream recovery location through the second cooling channel in the hub, thus establishing a rotor blade cooling circuit. Cooling air is drawn from a starting point in the preceding stage, passes sequentially through the first cooling channel in the casing, the cooling channel in the stator blades, and then returns to an upstream recovery location through the second cooling channel in the casing, thus establishing a stator blade cooling circuit. SS3. Based on the pressure field distribution within the compressor, a location with sufficient pressure to support the cooling airflow to complete the entire loop is selected as the initial starting position. A location where the recovered airflow pressure exceeds the local mainstream pressure is selected as the initial recovery position. Through fluid-structure interaction calculations, the starting and recovery positions are iteratively adjusted until the desired blade cooling amplitude and smooth cooling airflow recovery without backflow are achieved. SS4. A composite cooling structure consisting of heat pipes and a loop heat exchanger is installed in the rotor / stator blades, casing, and hub that require cooling. The heat pipes are embedded in the rotor and stator blades, respectively, with the evaporation ends of the heat pipes located in the high-heat-load area of ​​the blades. The loop heat exchangers are located in the hub and casing, respectively, with the hot side of the loop heat exchanger tightly connected to the condensation end of the heat pipe, and the cold side of the loop heat exchanger connected to the first cooling channel inside the hub and the first cooling channel inside the casing, respectively. SS5. Iteratively optimize the cooling solution through numerical simulation and / or experimental testing to evaluate the cooling design's effectiveness. Optimize and adjust the parameters of internal cooling channels, heat pipes, and / or loop heat exchangers to verify that the cooling design meets the requirements for blade cooling and smooth cooling airflow circulation, ultimately developing a final cooling solution.

2. The multi-stage axial flow compressor design method considering the final stage internal cooling according to claim 1 is characterized in that: In the above step SS1, the structure-aerodynamic coupling analysis model is three-dimensionally modeled based on the flow path and blade geometry of the multi-stage axial flow compressor, and the pressure field and temperature field distribution inside the compressor are solved using the RANS equations and turbulence model. The temperature rise on the surface and end area of ​​the compressor blade and the rate of turbulent pulsation generation are analyzed. At the same time, combined with the thermodynamic properties of the blade material, the number of stages and specific cooling areas that require cooling design are clearly determined.

3. The design method of a multi-stage axial flow compressor considering the final stage internal cooling according to claim 1 is characterized in that: In the above step SS2, in the rotor blade cooling circuit, the outlet position and recovery position of the cooling airflow are both located on the hub of the front stage of the compressor, and the outlet position is located at the high-pressure downstream stage, and the recovery position is located at the low-pressure upstream stage; in the stator blade cooling circuit, the outlet position and recovery position of the cooling airflow are both located on the casing of the front stage of the compressor, and the outlet position is located at the high-pressure downstream stage, and the recovery position is located at the low-pressure upstream stage.

4. The multi-stage axial flow compressor design method considering the final stage internal cooling according to claim 1 is characterized in that: In the above step SS3, during the optimization and determination of the starting and recovery positions of the cooling circuit, the cooling circuit layout is optimized using fluid-solid coupling iterative calculations. With the goal of minimizing cooling air flow and minimizing losses along the way, the path length, cross-sectional shape, and geometric dimensions of the cooling channel are optimized and adjusted, while meeting the requirements of the preset cooling amplitude of the blade and the smooth recovery of the cooling airflow without backflow.

5. The design method of a multi-stage axial flow compressor considering the final stage internal cooling according to claim 1, characterized in that: In the above step SS4, the heat pipe adopts a multi-point distributed arrangement scheme. According to the distribution of high heat load areas inside the last-stage blade, the evaporation end of the heat pipe is arranged at multiple points inside the blade, and uniform heat transfer is achieved through joint heat exchange between the heat pipes; at the same time, the condensation end of the heat pipe is tightly connected to the hot side of the loop heat exchanger using high thermal conductivity materials, and the layout of the condensation end is adapted to the geometric shape of the loop heat exchanger to improve the efficiency of heat transfer from the heat pipe to the cooling airflow.

6. The multi-stage axial flow compressor design method considering the final stage internal cooling according to claim 1 is characterized in that: In the above step SS5, the numerical simulation uses the unsteady RANS method to establish a full three-dimensional calculation model including blades, heat pipes and heat exchangers. The SST turbulence model based on the coupled high-precision transition format is used to accurately predict the heat transfer characteristics of the blade surface. The calculation results need to be iterated through fluid-heat coupling to obtain a converged solution for the blade surface temperature distribution and flow field characteristics.

7. The multi-stage axial flow compressor design method considering the final stage internal cooling according to claim 6 is characterized in that: When evaluating the impact of cooling design on the blade surface temperature distribution, boundary layer growth rate and end-zone turbulent pulsation, it mainly includes: obtaining the blade surface temperature distribution with and without cooling conditions through numerical simulation, focusing on the local temperature drop near the high heat load area of ​​the blade, and evaluating the cooling effect and uniformity of the cooling design; comparing the turbulent pulsation characteristics and boundary layer development laws of the suction surface of the blade mid-section with and without cooling conditions, and evaluating the role of the cooling effect in weakening the trailing edge separation of the blade mid-section and delaying the growth of the boundary layer; comparing the three-dimensional Reynolds normal stress at different axial positions near the end wall with and without cooling conditions, and evaluating the effect of the cooling effect in weakening the lateral / radial migration of the secondary flow in the end zone, reducing the three-dimensional mixing of turbulent flow and the total pressure loss of the outlet section.

8. The design method for a multi-stage axial flow compressor taking into account the final stage internal cooling according to claim 7, characterized in that: When iteratively optimizing the cooling scheme based on the evaluation and analysis results, the hub and casing cooling channels, heat pipe layout and loop heat exchanger design are adjusted to ensure the synergistic improvement of the cooling effect and the aerodynamic performance of the blade, including: optimizing the path length, cross-sectional shape and geometric dimensions of the cooling channel to achieve the minimum cooling air flow and the lowest along-the-path loss, and improving the flow uniformity; optimizing the material, diameter, length and layout of the heat pipe to improve the heat transfer efficiency and ensure the structural stability under high-speed rotation conditions; optimizing the heat transfer area, flow channel structure and installation position of the loop heat exchanger to achieve the maximum heat transfer and reduce the total pressure loss in a limited space; and the final cooling scheme should meet the requirements of the cooling amplitude of the blade, boundary layer growth rate control, average total pressure loss coefficient control of the outlet section, and smooth circulation of the cooling airflow.

9. A multi-stage axial flow compressor taking into account the final stage internal cooling, characterized in that: The design optimization is performed by adopting the multi-stage axial flow compressor design method considering the last stage internal cooling as described in any one of claims 1 to 8.

Images