Aero-engine turbine guide vane double-wall rim plate heat transfer analysis method

Abstract

An aero-engine turbine guide vane double-wall rim plate heat transfer analysis method, comprising: according to the pressure characteristic curve in the rim plate, the rim plate is divided into a plurality of calculation regions along the isobar gradient; in each calculation region, the film holes with the same order of magnitude of outlet back pressure are combined into a film hole unit, and the corresponding impingement holes are combined into an impingement hole unit, and the spoiler columns are combined into a spoiler column unit; the resistance elements and the heat exchange elements close to the real flow are extracted; the film hole unit, the impingement hole unit and the spoiler column unit in each calculation region are fitted in series with the resistance elements and the heat exchange elements to obtain a one-dimensional flow heat transfer calculation network; on the one-dimensional flow heat transfer calculation network, a refined flow heat transfer analysis based on the isobar is carried out to obtain the heat transfer coefficient and the heat transfer temperature of the cooling heat transfer surface in the rim plate; a three-dimensional heat conduction grid heat analysis model of the rim plate is constructed, the heat transfer coefficient and the heat transfer temperature are assigned on the heat transfer surface in each calculation region, and heat calculation is carried out to obtain a three-dimensional temperature field.

Description

Technical Field

[0001] This application belongs to the field of heat transfer analysis technology for double-walled edge plates of aero-engine turbine guide vanes, specifically relating to a heat transfer analysis method for double-walled edge plates of aero-engine turbine guide vanes. Background Technology

[0002] When an aero-engine is operating, the turbine guide vanes are directly subjected to the uneven high-temperature field at the combustion chamber outlet, and their rims are at extremely high risk of being burned.

[0003] To prevent the turbine guide vane rim from burning out, a double-walled structure is designed, forming a cavity between the outer and inner walls. Multiple impact holes are distributed on the outer wall, and multiple film cooling holes are distributed on the inner wall. Multiple flow-deflecting columns are installed within the cavity, each supported between the outer and inner walls. The outer and inner walls have perforations for the guide vane tip to insert into. During engine operation, cooling gas enters the cavity through the impact holes for impact cooling. After being turbulent by the flow-deflecting columns, the cooling gas exits through the film cooling holes, forming a film cooling system that ultimately mixes with the main combustion gas. This design provides highly efficient cooling and protection for the rim, effectively preventing rim burning.

[0004] The turbine guide vane double-walled edge plate integrates multiple cooling methods such as impingement jet, turbulence column flow, and film cooling. Its fluid flow and heat transfer relationship are complex. Accurately obtaining the three-dimensional temperature distribution of the edge plate and optimizing the temperature field uniformity is of great guiding significance for the design and improvement of the edge plate.

[0005] Currently, the three-dimensional temperature distribution of the blades is mostly obtained through special test runs of the double-walled blades of aero-engine turbines. This technical solution is time-consuming and labor-intensive, and due to space limitations, the temperature test points are limited to the sides of the gas inlet and outlet, and the number is limited. Especially for cases with a small number of guide vanes, low-consistency, long-chord blades, and uneven flow heat transfer distribution on the surface of the blades, it is difficult to accurately obtain the three-dimensional temperature distribution of the blades.

[0006] This application is made in view of the aforementioned technical deficiencies.

[0007] It should be noted that the above background information is only used to assist in understanding the inventive concept and technical solution of this invention, and it does not necessarily belong to the prior art of this patent application. In the absence of clear evidence that the above information was disclosed on the filing date of this application, the above background information should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0008] The purpose of this application is to provide a method for analyzing heat transfer in the double-walled edge plate of aero-engine turbine guide vanes, so as to overcome or mitigate at least one of the known technical defects.

[0009] The technical solution of this application is:

[0010] A method for analyzing heat transfer in a double-layered wall of an aero-engine turbine guide vane includes:

[0011] Based on the pressure characteristic curve inside the rim plate, the rim plate is divided into multiple calculation regions along the isobaric gradient.

[0012] Within each calculation region, film pores with outlet back pressure of the same order of magnitude are merged into film pore units, and impingement pores are merged into impingement pore units and turbulence columns are merged into turbulence column units.

[0013] Extract resistance elements and heat exchange elements that closely resemble real flow;

[0014] For the film pore unit, impact pore unit, and turbulence column unit of each calculation region, a one-dimensional flow heat transfer calculation network is obtained by fitting the resistance element and heat transfer element in series.

[0015] On a one-dimensional flow heat transfer calculation network, a refined flow heat transfer analysis based on isobars is carried out to obtain the heat transfer coefficient and heat transfer temperature of the cooling heat transfer surface inside the rim plate;

[0016] A three-dimensional thermally conductive mesh thermal analysis model of the shroud is constructed. The heat transfer coefficient and heat transfer temperature are assigned to the heat transfer surface of each computational region, and thermal calculations are performed to obtain the three-dimensional temperature field.

[0017] According to at least one embodiment of this application, in the above-described heat transfer analysis method for the double-walled rim of an aero-engine turbine guide vane, the rim is divided into multiple calculation regions along the isobar gradient based on the pressure characteristic curve inside the rim, specifically including the blade back calculation region, the blade base calculation region, the leading edge calculation region, and the trailing edge calculation region.

[0018] According to at least one embodiment of this application, in the above-described heat transfer analysis method for the double-walled edge plate of an aero-engine turbine guide vane, the extraction of resistance elements and heat transfer elements that closely approximate actual flow is specifically as follows:

[0019] Based on the aero-engine air system calculation program, and compared with the air system calculation module database, resistance elements and heat exchange elements that closely approximate real flow are selected.

[0020] According to at least one embodiment of this application, in the above-described heat transfer analysis method for the double-walled flange of an aero-engine turbine guide vane, the extracted resistance element and heat exchange element include:

[0021] The T13 element has heat exchange losses due to the array of circular jets impacting the flat wall.

[0022] T01 element, throttling orifice without considering flow loss and heat transfer;

[0023] The T08 element features a rectangular channel with the combined effects of transverse ribs and spoiler columns.

[0024] This application has at least the following beneficial technical effects:

[0025] A method for heat transfer analysis of the double-walled rim of aero-engine turbine guide vanes is provided. The method is designed to divide the rim into multiple computational regions along isobaric gradients based on the pressure characteristic curve within the rim. Within each computational region, film vents with outlet back pressure of the same order of magnitude are merged into film vent elements, and similarly, impingement vents and turbulence columns are merged into turbulence column elements. A one-dimensional flow heat transfer calculation network is obtained by fitting drag elements and heat transfer elements in series. A refined flow heat transfer analysis based on isobars is then performed to obtain the heat transfer coefficient and heat transfer temperature of the cooling heat transfer surface within the rim. Finally, in a three-dimensional thermally conductive mesh thermal analysis model, the heat transfer coefficient and heat transfer temperature are assigned to the heat transfer surface of each computational region for thermal calculation, resulting in a three-dimensional temperature field. The process is simple and has high accuracy. Attached Figure Description

[0026] Figure 1 This is a schematic diagram of the double-layer wall flange of the turbine guide vane provided in an embodiment of this application;

[0027] Figure 2 This is a schematic diagram of the internal pressure characteristic curve of the flange provided in the embodiment of this application;

[0028] Figure 3 This is a schematic diagram of the partitioning of the edge plate calculation region and the merging of its air film holes provided in the embodiments of this application;

[0029] Figure 4 This is a schematic diagram of the one-dimensional flow heat transfer calculation network of the rim plate provided in the embodiments of this application;

[0030] Figure 5 This is a schematic diagram of the heat transfer zone for calculating the three-dimensional temperature field of the shroud provided in the embodiments of this application.

[0031] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual size of the product. Furthermore, the accompanying drawings are for illustrative purposes only and should not be construed as limiting this patent. Detailed Implementation

[0032] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings. Other related parts can be referred to the general design. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined with each other to obtain new embodiments.

[0033] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer," etc., used in this application description to indicate relative direction or positional relationship are used only to indicate relative orientation or positional relationship, and do not imply that the device or component must have a specific orientation, or be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, and therefore should not be construed as a limitation on this application. The terms "first," "second," "third," and similar terms used in this application description are used only for descriptive purposes to distinguish different components, and should not be construed as indicating or implying relative importance. The terms "a," "one," or "the," etc., used in this application description should not be construed as an absolute limitation on quantity, but should be construed as indicating the existence of at least one. The terms "including," "comprising," etc., used in this application description mean that the element or object preceding the word covers the element or object listed after the word and its equivalents, without excluding other elements or objects.

[0034] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as “installation,” “connection,” and “linkage” used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium; or it can be a connection within two components. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0035] The following is in conjunction with the appendix Figures 1 to 5 This application will be described in further detail.

[0036] When analyzing the heat transfer of the double-walled rim plate of an aero-engine turbine guide vane, it is necessary to understand the flow characteristics within the rim plate. The rim plate structure has the following unique properties:

[0037] a) The interlayer cavity is a large through-hole cooling chamber. The cooling gas flows out from the impact hole and impacts the gas film hole surface. It flows freely in the interlayer cavity. The flow state of the cooling gas depends entirely on the relative back pressure at the outlet of each exhaust film hole. From this perspective, the fluid flow in the interlayer cavity can be regarded as unrestrained and uncontrolled.

[0038] b) Back pressure distribution of the main combustion gas flow channel at the outlet of the gas film holes at various positions on the flange, that is, the pressure and temperature distribution of the main combustion fluid on the flange flow channel surface;

[0039] c) The overall profile of the rim plate is close to a slender rhombus, and the rim plate of the guide vane with large chord length and low density is even more slender;

[0040] The unique structure of the double-layered rim plate of the turbine guide vane of an aero-engine results in extremely drastic pressure changes within the rim plate channels, allowing for calculation and partitioning of the rim plate based on these pressure variations.

[0041] Pressure characteristic curve inside the flange, such as Figure 2 As shown, the calculation regions are divided into multiple regions based on the isobaric gradient. The calculation regions are the basis of the calculation and directly determine the accuracy of the temperature field calculation. They can be divided into the blade back calculation region, blade basin calculation region, leading edge calculation region, and trailing edge calculation region.

[0042] Without compromising the accuracy of engineering calculations, the film cooling holes within each calculation region are grouped and combined into film cooling hole units. The boundary of the film cooling holes in the central region of the combination represents the outflow boundary of the combined holes. Furthermore, the impinging holes within these units are combined into impinging hole units, and the turbulence columns are combined into turbulence column units. This significantly reduces computational costs. The rationality of the film cooling hole combination directly affects the accuracy of the rim cooling air consumption calculation. See details... Figure 3 .

[0043] For the blade back calculation region, the pressure variation at the outlet of the film air vents on the blade back is large in the chord region. The blade back calculation region can be refined into three heat transfer characteristic zones: SB1, SB2, and SB3. Each zone is characterized by a large pressure variation gradient along the normal of the blade-shaped vents. The film air vents are grouped in the direction perpendicular to the wall.

[0044] For the leaf-pot calculation region SP1, the isobars are almost perpendicular to the leaf-shaped aperture normal, and the air film apertures are grouped along the chord direction.

[0045] For the leading edge calculation region, judging from the pressure distribution cloud map inside the blade, it is a high-pressure zone. The air film pores with the same back pressure are divided into a region, and the corresponding heat transfer and impact heat transfer regions of the turbulence column region are defined accordingly. Specifically, it can be subdivided into the leading edge blade back calculation region SQB, the leading edge blade basin calculation region SQP, and the positive leading edge calculation region SQ.

[0046] For the SQB calculation region on the back of the blade, the pressure change in the flow channel is most drastic due to airflow disturbance. When grouping and merging the air film in this region, the outflow area of ​​the outlet orifice unit must be on the same order of magnitude, and the outlet back pressure must also be on the same order of magnitude. This region occupies one diagonal of the rhomboid flange, making the heat transfer in this partition extremely irregular. When extracting the columnar geometric characteristic parameters for calculating the heat transfer characteristics, it is necessary to consider converting them into corresponding regular columnar characteristic parameters.

[0047] For the leading edge leaf basin calculation region SQP, the pressure change is gentler than that of the leaf back calculation region SQB, and the corresponding solid area is larger. The isobars are perpendicular to the leaf shape aperture normal and there are more air film pores in the direction perpendicular to the wall. The air film pores are grouped along the chord direction.

[0048] For the positive leading edge calculation region SQ, the pressure is very high, and the film pores in this region can be merged.

[0049] For the trailing edge calculation region, it can be refined into four heat transfer characteristic zones: SWP, SW1, SW2, and SW3. The characteristic parameters extracted from the heat transfer in the turbulence column region within SW3 are along the wall normal. The pressure changes drastically and include an extremely low pressure region in the area from the middle chord of the guide vane back to the trailing edge outlet. The isopressure distribution curves are almost parallel to the contour lines at other locations in the channel, but here they are closed vortex lines. The characteristic parameters of the other three heat transfer characteristic zones, SWP, SW1, and SW2, are extracted along the chord direction.

[0050] Based on the aero-engine air system calculation program, and by comparing with the air system calculation module database, drag elements and flow heat transfer elements that closely approximate real flow are extracted.

[0051] Based on the aero-engine air system calculation program, and comparing it with the air system calculation module database, drag and heat transfer elements that closely resemble real flow are selected. For the film atomizer elements, impingement atomizer elements, and turbulence column elements in each calculation region, a one-dimensional flow and heat transfer calculation network is obtained by fitting drag and heat transfer elements in series. This simplifies the complex three-dimensional flow of the rim plate into a one-dimensional network topology, such as... Figure 4 As shown.

[0052] The extracted resistance elements and heat exchange elements include:

[0053] The T13 element has an array-type circular jet impacting flat wall loss for heat exchange, which is used to model the impact heat transfer.

[0054] The T01 element is a throttling orifice that does not consider flow losses and heat transfer, used to model the gas film outflow.

[0055] The T08 element features a rectangular channel with the combined effects of transverse ribs and turbulence pillars, used to model turbulent flow.

[0056] While transforming three-dimensional flow into one-dimensional flow, the process eliminates factors irrelevant to heat transfer calculations, ensuring that the simplified process retains the characteristic parameters of the double-walled edge plate, including impact, turbulence, and film outflow.

[0057] Based on a one-dimensional network, a refined flow and heat transfer analysis based on the isobars of the flange is carried out. Aerodynamic boundary parameters, heat transfer coefficient and temperature of the gas-fuel side of the flange, resistance elements, geometric parameters required by heat transfer elements, cooling gas supply pressure and temperature, and the back pressure of the gas film orifice outflow are the combined outflow pressure of the gas film orifice in each zone. Finally, the cooling air consumption of the flange, the heat transfer coefficient of the internal cooling heat transfer surface, and the heat transfer temperature are obtained.

[0058] The rim plate is meshed into a solid model in three dimensions, such as... Figure 5 The mesh cells are partitioned as shown, corresponding to each computational region. The heat transfer coefficient and temperature of the third type of boundary condition are assigned to the heat transfer surface corresponding to each partition. Based on the thermal analysis module of ANSYS analysis software, the rim plate is calculated to obtain the three-dimensional temperature field of the rim plate.

[0059] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of this application.

Claims

1. A method for analyzing heat transfer in a double-layered wall of a turbine guide vane of an aero-engine, characterized in that, include: Based on the pressure characteristic curve inside the rim plate, the rim plate is divided into multiple calculation regions along the isobaric gradient. Within each calculation region, film pores with outlet back pressure of the same order of magnitude are merged into film pore units, and impingement pores are merged into impingement pore units and turbulence columns are merged into turbulence column units. Extract the resistance elements and heat exchange elements of the actual flow; For the film pore unit, impact pore unit, and turbulence column unit of each calculation region, a one-dimensional flow heat transfer calculation network is obtained by fitting the resistance element and heat transfer element in series. On a one-dimensional flow heat transfer calculation network, a refined flow heat transfer analysis based on isobars is carried out to obtain the heat transfer coefficient and heat transfer temperature of the cooling heat transfer surface inside the rim plate; A three-dimensional thermally conductive mesh thermal analysis model of the shroud is constructed. The heat transfer coefficient and heat transfer temperature are assigned to the heat transfer surface of each computational region, and thermal calculations are performed to obtain the three-dimensional temperature field.

2. The heat transfer analysis method for the double-layer wall flange of an aero-engine turbine guide vane according to claim 1, characterized in that, The blade is divided into multiple calculation regions along the isobar gradient based on the pressure characteristic curve inside the blade, specifically including the blade back calculation region, blade base calculation region, leading edge calculation region, and trailing edge calculation region.

3. The heat transfer analysis method for the double-layer wall flange of an aero-engine turbine guide vane according to claim 2, characterized in that, Extracting the resistance elements and heat exchange elements of the actual flow, specifically: Based on the aero-engine air system calculation program, and by comparing with the air system calculation module database, the drag elements and heat exchange elements of the actual flow are extracted.

4. The heat transfer analysis method for the double-layer wall flange of an aero-engine turbine guide vane according to claim 1, characterized in that, The extracted resistance elements and heat exchange elements include: The T13 element has heat exchange losses due to the array of circular jets impacting the flat wall. T01 element, throttling orifice without considering flow loss and heat transfer; The T08 element features a rectangular channel with the combined effects of transverse ribs and spoiler columns.

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