A method for predicting axial force on aero-engine rotors

Abstract

This application belongs to the field of aero-engine design technology, specifically involving a method for predicting the axial force of an aero-engine rotor. Based on the parameters provided by the aero-engine aerodynamic scheme, the method constructs a simplified model of the turbine blade cooling flow characteristics, which can quickly and accurately predict and adjust the rotor axial force during the aero-engine scheme design stage, reduce the iteration of aero-engine design, and promote the development of aero-engines.

Description

Technical Field

[0001] This application belongs to the field of aero-engine design technology, specifically relating to a method for predicting the axial force of an aero-engine rotor. Background Technology

[0002] The axial force of an aero-engine rotor refers to the aerodynamic load acting on the rotor, which is transmitted to the load-bearing frame through the bearings supporting the rotor.

[0003] The airflow channels of an aero-engine include the main flow path and the air system flow path. The air system flow path is a secondary flow path, and its airflow mainly comes from the compressor. According to the required pressure, temperature and flow rate, the airflow is led out at an appropriate location and flows through structures such as disc cavity, vent hole, sealing device, and pipeline to cool high-temperature components, seal the bearing cavity and high-temperature main flow path, and control the axial load of the bearing.

[0004] The axial force of an aero-engine rotor mainly includes the main flow path axial force and the disk cavity axial force. The disk cavity axial force is derived based on the air system flow path calculation and is affected by the turbine blade cooling design.

[0005] The axial force of an aero-engine rotor changes with its operating conditions. During the design phase, it is necessary not only to ensure that the rotor axial force meets the requirements under the design conditions, but also to ensure that the rotor axial force meets the requirements under each operating condition within the flight envelope. The rotor axial force should not be too large to avoid exceeding the bearing's load-bearing capacity, nor should it be too small to avoid causing bearing slippage damage.

[0006] Currently, the assessment of axial force on aero-engine rotors is often conducted after the turbine blade cooling scheme design is completed. If the assessment does not meet the requirements, the required iteration cycle is long, and it may even require a complete overhaul of the overall structural design, which seriously restricts the design process of aero-engines.

[0007] This application is made in view of the aforementioned technical deficiencies.

[0008] 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 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

[0009] The purpose of this application is to provide a method for predicting the axial force of an aero-engine rotor, so as to overcome or mitigate at least one of the known technical defects.

[0010] The technical solution of this application is:

[0011] A method for predicting the axial force of an aero-engine rotor includes:

[0012] Determine the cooling scheme for each row of turbine cooling blades and obtain the cooling effect of each row of turbine cooling blades;

[0013] Based on the cooling scheme and cooling effect of each row of turbine cooling blades, check the cooling air volume of each row of turbine cooling blades.

[0014] Design the overall structural scheme and determine the dimensions of the air system flow path structure;

[0015] Identify the turbine cooling blades that affect the rotor's axial force, and determine the temperature and pressure of the cooling gas at the inlet and outlet of these turbine cooling blades;

[0016] A simplified model of the cooling air flow rate of the turbine cooling blades that affects the rotor axial force is constructed. Under the condition that the temperature and pressure of the inlet and outlet cooling air, as well as the aerodynamic parameters of the main flow path are determined, the cooling air flow rate is adjusted to match the verified cooling air flow rate.

[0017] Based on the air system flow path structure dimensions and a simplified model of the cooling air flow rate of the turbine cooling blades that affect the rotor axial force, the air system fluid dynamics calculations are carried out to obtain the pressure of each disk cavity;

[0018] Using the main flow aerodynamic parameters, calculate the main flow axial force of the rotor compression system components and turbine components. Using the pressure of each disk cavity, calculate the disk cavity axial force of the rotor compression system components and turbine components. Superimpose the main flow axial force and disk cavity axial force of the rotor compression system components and turbine components to obtain the rotor axial force.

[0019] When the rotor axial force does not meet the requirements, the air system flow path structure dimensions are readjusted until the rotor axial force meets the requirements, satisfying the design requirements or reaching the limit of the disk cavity axial force adjustment.

[0020] According to at least one embodiment of this application, in the above-described method for predicting the axial force of an aero-engine rotor, the determination of the cooling scheme for each row of turbine cooling blades is specifically made based on the overall requirements, aerodynamic scheme, and material selection scheme of the aero-engine, combined with the technical level and experience of turbine blade cooling design.

[0021] According to at least one embodiment of this application, in the above-described method for predicting the axial force of an aero-engine rotor, obtaining the cooling effect of each row of turbine cooling blades specifically involves:

[0022]

[0023] in,

[0024] η represents the cooling effect of the turbine cooling blades;

[0025] T g The average temperature of the combustion gas inlet to the turbine cooling blades;

[0026] T b The average temperature of the turbine blades;

[0027] T c This represents the average temperature of the cooling air for the turbine blades.

[0028] According to at least one embodiment of this application, in the above-described method for predicting the axial force of an aero-engine rotor, the step of checking the cooling air volume of each row of turbine cooling blades based on the cooling scheme and cooling effect of each row of turbine cooling blades specifically involves checking the cooling effect of the turbine cooling blades' cooling scheme and the percentage relationship between the cooling air volume and the compressor flow rate.

[0029] According to at least one embodiment of this application, in the above-described method for predicting the axial force of an aero-engine rotor, the step of designing an overall structural scheme and determining the dimensions of the air system flow path structure specifically involves designing an overall structural scheme based on functional requirements, and then determining the dimensions of the air system flow path structure.

[0030] According to at least one embodiment of this application, in the above-described method for predicting the axial force of an aero-engine rotor, the step of determining the turbine cooling blades that affect the rotor axial force, and determining the temperature and pressure of the inlet and outlet cooling gases of the turbine cooling blades, specifically involves:

[0031] Based on the main aerodynamic parameters, air system flow path characteristics, structural dimensions, throttling units, and relevant experience with turbine cooling blades, the turbine cooling blades that affect the axial force of the engine rotor are determined, and then the temperature and pressure of the inlet and outlet cooling air are determined.

[0032] According to at least one embodiment of this application, in the above-described method for predicting the axial force of an aero-engine rotor, the turbine cooling blades that affect the rotor axial force include high-pressure turbine rotor blades, low-pressure turbine first-stage guide vanes, and low-pressure turbine blades that employ cooling design and affect the rotor axial force.

[0033] According to at least one embodiment of this application, in the above-described method for predicting the axial force of an aero-engine rotor, the step of constructing a simplified model of the cooling air flow rate of the turbine cooling blades that affects the rotor axial force specifically involves:

[0034] For turbine rotor blades that employ composite cooling of film, impact, and convection, an inlet cavity unit and an inlet hole unit are provided corresponding to the cooling gas inlet, an inner cavity unit is provided corresponding to the inlet cavity, and a leading edge hole unit, blade basin hole unit, blade back hole unit, blade tip hole unit, and tail slot unit are provided corresponding to the cooling gas outlet. Furthermore, a correction hole unit and a transition cavity unit are added between the inner cavity unit and the tail slot unit.

[0035] For turbine rotor blades employing divergent or quasi-divergent cooling schemes, an inlet cavity unit and an inlet hole unit are provided corresponding to the cooling gas inlet; an inner cavity unit is provided corresponding to the inlet chamber; a leading edge inner hole unit, blade basin inner hole unit, and blade back inner hole unit are provided corresponding to the impact hole; a leading edge outer hole unit, blade basin outer hole unit, blade back outer hole unit, blade tip hole unit, and tail slot unit are provided corresponding to the cooling gas outlet; and an interlayer cavity unit is provided between the leading edge inner hole unit and the leading edge outer hole unit, between the blade basin inner hole unit and the blade basin outer hole unit, and between the blade back inner hole unit and the blade back outer hole unit corresponding to the impact chamber; and a correction hole unit and a transition cavity unit are added between the inner cavity unit and the tail slot unit.

[0036] For the first-stage guide vane of a low-pressure turbine with a composite cooling design, an inlet cavity unit and an inlet hole unit are set corresponding to the cooling gas inlet, an inner cavity unit is set corresponding to the inlet chamber, and a leading edge hole unit, blade basin hole unit, blade back hole unit, outlet hole unit, and tail slot unit are set corresponding to the cooling gas outlet. A correction hole unit and a transition cavity unit are added between the inner cavity unit and the tail slot unit.

[0037] This application has at least the following beneficial technical effects:

[0038] This paper presents a method for predicting the axial force of an aero-engine rotor. Based on the parameters provided by the aero-engine aerodynamic scheme, a simplified model of the turbine blade cooling flow characteristics is constructed. This method can quickly and accurately predict and adjust the rotor axial force during the aero-engine scheme design stage, reduce the iteration of aero-engine design, and promote the development of aero-engines. Attached Figure Description

[0039] Figure 1 This is a schematic diagram of the method for predicting the axial force of an aero-engine rotor provided in an embodiment of this application;

[0040] Figure 2 This is a diagram showing the cooling effect of the turbine cooling blade cooling scheme provided in this application embodiment, and the percentage relationship between the cooling gas and the compressor flow rate;

[0041] Figure 3 This is a schematic diagram of a simplified model of cooling airflow constructed using a composite cooling turbine rotor blade employing film, impact, and convection cooling, provided in an embodiment of this application.

[0042] Figure 4 This is a schematic diagram of a simplified model of cooling airflow for turbine rotor blades employing a divergent or quasi-divergent cooling scheme, provided in an embodiment of this application.

[0043] Figure 5 This is a schematic diagram of a simplified model of cooling airflow constructed using a composite cooling design for the first-stage guide vanes of a low-pressure turbine, as provided in an embodiment of this application.

[0044] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual product dimensions. Furthermore, the drawings are for illustrative purposes only and should not be construed as limiting this application. Detailed Implementation

[0045] 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.

[0046] 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.

[0047] 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.

[0048] The following is in conjunction with the appendix Figures 1 to 5The method for predicting the axial force of an aero-engine rotor provided in this application is described in further detail.

[0049] Step 1: Based on the overall requirements, aerodynamic scheme, and material selection scheme of the aero-engine, and combined with the technical level and experience of turbine blade cooling design, determine the cooling scheme for each row of turbine cooling blades, and obtain the cooling effect of each row of turbine cooling blades:

[0050]

[0051] in,

[0052] η represents the cooling effect of the turbine cooling blades;

[0053] T g The average temperature of the combustion gas inlet to the turbine cooling blades;

[0054] T b The average temperature of the turbine blades;

[0055] T c This represents the average temperature of the cooling air for the turbine blades.

[0056] Step 2: Based on the cooling scheme and cooling effect of each row of turbine cooling blades, verify the cooling air volume of each row of turbine cooling blades. (See [reference]) Figure 2 .

[0057] Step 3: Based on functional requirements, design the overall structural scheme and determine the main flow path and structural dimensions of the air system.

[0058] Step 4: Based on the engine's main aerodynamic parameters, air system flow path characteristics, structural dimensions, throttling unit, and relevant experience with turbine cooling blades, determine the pressure and temperature of the inlet and outlet cooling air of the turbine cooling blades that affect the axial force of the engine rotor.

[0059] The turbine cooling blades that affect the axial force of the engine rotor mainly include high-pressure turbine rotor blades, low-pressure turbine first-stage guide vanes, and other low-pressure turbine blades that adopt cooling design and affect the rotor axial force. Among them, the cooling gas inlet of high-pressure turbine rotor blades and low-pressure turbine rotor blades is generally located at the bottom of the tenon, and the cooling gas outlet is located at the blade body and blade tip; the cooling gas inlet of low-pressure turbine first-stage guide vanes is generally located on the outer side of the upper edge plate, and the cooling gas outlet is located on the outer side of the lower edge plate and blade body.

[0060] Step 5: Based on the aerodynamic parameters of the main flow path and the outer wall surface of the high-pressure turbine rotor blades, low-pressure turbine rotor blades, and low-pressure turbine guide vanes, and combined with the cooling flow required at different outlet positions of the cooling gas, construct a simplified model of the cooling flow characteristics of the high-pressure turbine rotor blades, low-pressure turbine rotor blades, and low-pressure turbine first-stage guide vanes.

[0061] A simplified model of the cooling flow characteristics of turbine rotor blades employing a combination of film cooling, impingement cooling, and convection cooling is constructed as follows: Figure 3 As shown, the cooling gas inlet is located below the tenon connecting the turbine rotor blade and the disk. The cooling gas outlets are located at the leading edge, blade basin side, blade back side, trailing edge, and blade tip. The size and number of the leading edge hole, blade basin hole, blade back hole, blade tip hole, and trailing edge slot are set according to the temperature of the gas flow in each part and the proportion of cooling gas distribution. They are also adjusted appropriately to ensure that the cooling gas flow rate matches the verified cooling gas flow rate. For a simple convection model, the outlet position of the cooling gas flow is appropriately selected and the cooling gas flow rate is adjusted based on the simplified model of the cooling flow characteristics of the composite cooling turbine rotor blade.

[0062] For turbine rotor blades employing divergent or quasi-divergent cooling schemes, a simplified model of the cooling flow characteristics is constructed as follows: Figure 4 As shown. The cooling gas inlet is located below the tenon connecting the turbine rotor blade and the disk. The cooling gas outlets are located at the leading edge, blade head, blade back, trailing edge, and blade tip. The blade profile adopts a double-wall structure. Inner holes are added to the inner wall for the leading edge, blade back, and blade head, respectively, to provide impact cooling to the outer wall, increasing the heat exchange area and improving the cooling effect. Cooling gas outlets / tail slots are provided on the outer wall for the leading edge, blade head, blade back, and trailing edge. The size and number of inner holes, outer holes, tail slots, and blade tip holes are set according to the gas flow temperature and cooling gas distribution ratio at each location, and are adjusted appropriately to ensure that the turbine blade cooling flow rate matches the verified cooling gas flow rate.

[0063] Low-pressure turbine first-stage guide vanes require both internal cooling design and often function as cooling airflow channels to introduce cooling airflow into the engine disk cavity. A simplified model of the cooling flow characteristics of low-pressure turbine first-stage guide vanes employing a composite cooling design is constructed as follows: Figure 5 As shown, the cooling gas inlet is located on the outer side of the upper edge plate, and the cooling gas outlets are located at the leading edge, blade head, blade back, and trailing edge of the blade body. Based on the temperature of the gas flow in each part and the proportion of cooling gas distribution, the size and number of the leading edge hole, blade head hole, blade back hole, and trailing slot are set and adjusted appropriately to ensure that the cooling flow rate of the first-stage guide vane of the low-pressure turbine and the cooling gas introduced into the disk cavity from the outlet cavity match the checked cooling gas flow rate. For low-pressure turbine guide vanes that only require simple convection cooling design, the outlet position of the cooling gas flow can be appropriately selected and the cooling gas flow rate adjusted based on the simplified model of the cooling flow characteristics of the composite cooling low-pressure turbine first-stage guide vane.

[0064] Step 6: Based on the size and characteristics of each throttling unit in the air system flow path and the simplified model of the turbine blade cooling flow characteristics, perform air system hydrodynamic calculations to obtain the pressure of each disk cavity that affects the rotor axial force.

[0065] Step 7: Using the aerodynamic parameters of the inlet and outlet of the main flow rotor blades of the compression system components and turbine components, including pressure, velocity, flow rate, and structural parameters, as input, calculate the rotor axial force of the main flow of the compression system components and turbine components using the developed program. At the same time, using the dimensions and pressure of each disk cavity as input files, calculate the axial force of the disk cavities of the compression system components and turbine components using the developed program. Superimpose the main flow axial force and the disk cavity axial force to calculate the rotor axial force.

[0066] If the rotor axial force does not meet the requirements, adjust the air system flow path and related structural dimensions, including the type, number and structural dimensions of the throttling unit, and the dimensions of the disk cavity structure, until the rotor axial force meets the requirements, satisfies the design requirements, or reaches the adjustment limit of the disk cavity axial force.

[0067] The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to each other.

[0068] 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 predicting the axial force of an aero-engine rotor, characterized in that, include: Determine the cooling scheme for each row of turbine cooling blades and obtain the cooling effect of each row of turbine cooling blades; Based on the cooling scheme and cooling effect of each row of turbine cooling blades, check the cooling air volume of each row of turbine cooling blades. Design the overall structural scheme and determine the dimensions of the air system flow path structure; Identify the turbine cooling blades that affect the rotor's axial force, and determine the temperature and pressure of the cooling gas at the inlet and outlet of these turbine cooling blades; A simplified model of the cooling air flow rate of the turbine cooling blades that affects the rotor axial force is constructed. Under the condition that the temperature and pressure of the inlet and outlet cooling air, as well as the aerodynamic parameters of the main flow path are determined, the cooling air flow rate is adjusted to match the verified cooling air flow rate. Based on the air system flow path structure dimensions and a simplified model of the cooling air flow rate of the turbine cooling blades that affect the rotor axial force, the air system fluid dynamics calculations are carried out to obtain the pressure of each disk cavity; Using the main flow aerodynamic parameters, calculate the main flow axial force of the rotor compression system components and turbine components. Using the pressure of each disk cavity, calculate the disk cavity axial force of the rotor compression system components and turbine components. Superimpose the main flow axial force and disk cavity axial force of the rotor compression system components and turbine components to obtain the rotor axial force. If the rotor axial force does not meet the requirements, readjust the air system flow path structure dimensions until the rotor axial force meets the requirements.

2. The method for predicting the axial force of an aero-engine rotor according to claim 1, characterized in that, The determination of the cooling scheme for each row of turbine cooling blades is specifically based on the overall requirements of the aero-engine, the aerodynamic scheme, the material selection scheme, and the technical level and experience of turbine blade cooling design.

3. The method for predicting the axial force of an aero-engine rotor according to claim 1, characterized in that, The cooling effect of each row of turbine cooling blades is obtained as follows: in, η represents the cooling effect of the turbine cooling blades; T g The average temperature of the combustion gas inlet to the turbine cooling blades; T b The average temperature of the turbine blades; T c This represents the average temperature of the cooling air for the turbine blades.

4. The method for predicting the axial force of an aero-engine rotor according to claim 1, characterized in that, The cooling air volume of each row of turbine cooling blades is checked based on the cooling scheme and cooling effect of each row of turbine cooling blades. Specifically, the check is performed based on the cooling effect of the turbine cooling blade cooling scheme and the percentage relationship between the cooling air and the compressor flow rate.

5. The method for predicting the axial force of an aero-engine rotor according to claim 1, characterized in that, The overall structural design involves determining the dimensions of the air system flow path structure based on functional requirements.

6. The method for predicting the axial force of an aero-engine rotor according to claim 1, characterized in that, The determination of the turbine cooling blades that affect the rotor axial force, specifically the determination of the temperature and pressure of the cooling gas at the inlet and outlet of the turbine cooling blades, is as follows: Based on the main aerodynamic parameters, air system flow path characteristics, structural dimensions, throttling units, and relevant experience with turbine cooling blades, the turbine cooling blades that affect the axial force of the engine rotor are determined, and then the temperature and pressure of the inlet and outlet cooling air are determined.

7. The method for predicting the axial force of an aero-engine rotor according to claim 1, characterized in that, Turbine cooling blades that affect the rotor axial force include high-pressure turbine rotor blades, low-pressure turbine first-stage guide vanes, and low-pressure turbine blades that employ cooling design and affect the rotor axial force.

8. The method for predicting the axial force of an aero-engine rotor according to claim 1, characterized in that, The simplified model for the cooling gas flow rate of the turbine cooling blades that affects the rotor axial force is constructed as follows: For turbine rotor blades that employ composite cooling of film, impact, and convection, an inlet cavity unit and an inlet hole unit are provided corresponding to the cooling gas inlet, an inner cavity unit is provided corresponding to the inlet cavity, and a leading edge hole unit, blade basin hole unit, blade back hole unit, blade tip hole unit, and tail slot unit are provided corresponding to the cooling gas outlet. Furthermore, a correction hole unit and a transition cavity unit are added between the inner cavity unit and the tail slot unit. For turbine rotor blades employing divergent or quasi-divergent cooling schemes, an inlet cavity unit and an inlet hole unit are provided corresponding to the cooling gas inlet; an inner cavity unit is provided corresponding to the inlet chamber; a leading edge inner hole unit, blade basin inner hole unit, and blade back inner hole unit are provided corresponding to the impact hole; a leading edge outer hole unit, blade basin outer hole unit, blade back outer hole unit, blade tip hole unit, and tail slot unit are provided corresponding to the cooling gas outlet; and an interlayer cavity unit is provided between the leading edge inner hole unit and the leading edge outer hole unit, between the blade basin inner hole unit and the blade basin outer hole unit, and between the blade back inner hole unit and the blade back outer hole unit corresponding to the impact chamber; and a correction hole unit and a transition cavity unit are added between the inner cavity unit and the tail slot unit. For the first-stage guide vane of a low-pressure turbine with a composite cooling design, an inlet cavity unit and an inlet hole unit are set corresponding to the cooling gas inlet, an inner cavity unit is set corresponding to the inlet chamber, and a leading edge hole unit, blade basin hole unit, blade back hole unit, outlet hole unit, and tail slot unit are set corresponding to the cooling gas outlet. A correction hole unit and a transition cavity unit are added between the inner cavity unit and the tail slot unit.

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