Variable air volume system and control method thereof

Abstract

This invention relates to the field of aero-engine technology and discloses a variable flow air system and its control method. Cooling airflow is drawn from the intermediate stage of the compressor through a first bleed pipe and a second bleed pipe. The cooling airflow from the first bleed pipe passes through the inner cavity of the first-stage guide vane of the low-pressure turbine, then exits through a first exhaust port and a first pre-swirl nozzle before entering the main flow path. The cooling airflow from the second bleed pipe, after passing through a variable cooling air control device, sequentially passes through the inner cavity of the second-stage guide vane of the low-pressure turbine, the front rim jet hole, and the rear rim jet hole before exiting and entering the main flow path. This invention can continue to cool components such as the inner support ring even when the cooling airflow changes or the pressure inside the disk cavity decreases, preventing component overheating and failure, maintaining the cooling function of the air system components, and solving the problems of blade backflow, large changes in rotor axial force, and overheating of rim components during variable flow processes.

Description

Technical Field

[0001] This invention relates to the field of aero-engine technology and discloses a variable flow air system and its control method. Background Technology

[0002] The air system of an aero-engine is closely related to the working environment and overall performance of engine components, and is one of the most important systems of the engine. It typically draws air from appropriate locations in the compression system, designs various flow units to force the gas to flow in the expected direction and with the specified airflow parameters to perform its designated functions, and finally discharges it from certain designated locations, either merging with the main flow or leaking directly to the outside of the engine. With the continuous improvement of aero-engine performance and the increasing demands for lifespan and reliability, turbine inlet temperatures are rising, requiring larger volumes of bleed air from the air system.

[0003] In recent years, the total bleed air volume of advanced high-performance engines has reached about 30% of the compressor inlet flow rate. However, this bleed air volume is determined based on the cooling requirements of the engine under maximum thermal load conditions. Under lower thermal load conditions such as engine subsonic cruise, the total bleed air volume is seriously excessive, which greatly affects the overall cycle efficiency of the engine and leads to higher fuel consumption.

[0004] Domestic and international scholars, drawing on the concept of variable-cycle engine design, have proposed various variable-flow air system design ideas to reduce bleed air volume and improve overall engine performance during subsonic cruise. However, these design ideas all suffer from irreconcilable design contradictions, preventing their implementation in engine designs. The main reason for these design contradictions is that as the cooling airflow rate decreases, its pressure decreases simultaneously, leading to air system malfunction. This malfunction manifests in three main ways:

[0005] 1) Reducing the cooling airflow rate leads to a decrease in the cooling airflow pressure on the turbine blades, preventing the airflow from the turbine blade film vents from flowing out normally, resulting in backflow of combustion gas and failure of blade cooling.

[0006] 2) Reducing the cooling airflow leads to a decrease in the cooling airflow pressure in the disk cavity, causing a significant change in the rotor axial force and rendering the axial force adjustment function ineffective;

[0007] 3) Reducing the cooling airflow leads to a decrease in the flange sealing pressure, resulting in backflow of gas into the rotor flange, overheating of components, and failure of component cooling function.

[0008] Research on variable flow rate (VFR) air systems is still in its early stages. Based on available information, the main VFR system disclosed domestically and internationally is the high-pressure turbine guide vane flow path of the AL-31F engine, which proposes a method to increase the cooling air flow at maximum engine speed. However, this method only adjusts the guide vane flow and cannot improve the overall engine cycle efficiency. Therefore, existing air system designs lack practically applicable VFR systems that can reduce cruise fuel consumption. Summary of the Invention

[0009] The purpose of this invention is to provide a variable flow air system and its control method, which can solve the problems of blade gas backflow, large changes in rotor axial force, and overheating of wheel rim components during variable flow processes.

[0010] To achieve the above-mentioned technical effects, the technical solution adopted by the present invention is as follows:

[0011] A variable flow air system, comprising:

[0012] The first air intake pipe has its inlet end connected to the intermediate stage of the compressor and its outlet end connected to the inner cavity of the first stage guide vane of the low-pressure turbine.

[0013] The first exhaust port is provided on the inner support ring of the first stage of the low-pressure turbine, and the outlet end of the first exhaust port faces the upstream high-pressure turbine rotor. The inner support ring of the first stage of the low-pressure turbine is also provided with a first pre-swirl nozzle facing the downstream low-pressure turbine first stage rotor.

[0014] The second air intake pipe has its inlet end connected to the intermediate stage of the compressor and its outlet end connected to the inner cavity of the guide vane of the second stage of the low-pressure turbine.

[0015] The front rim jet hole is provided on the inner support ring of the second stage of the low-pressure turbine, and the outlet end of the front rim jet hole faces the rim of the first stage rotor of the low-pressure turbine upstream.

[0016] The rear rim jet hole is provided on the inner support ring of the second stage of the low-pressure turbine, and the outlet end of the rear rim jet hole faces the rim of the downstream second stage rotor of the low-pressure turbine.

[0017] A variable cooling air control device is installed on the second air intake pipe to adjust the cooling airflow rate of the second air intake pipe.

[0018] Furthermore, the cooling film holes of the first-stage guide vane of the low-pressure turbine are arranged on the blade basin side and the blade back cheek area, and there is no cooling airflow at the leading edge and trailing edge of the first-stage guide vane of the low-pressure turbine.

[0019] Furthermore, it also includes a shared disk cavity, which is formed by the low-pressure turbine second-stage inner support ring, the low-pressure turbine first-stage rotor, and the low-pressure turbine second-stage rotor; a second pre-swirl nozzle is also provided on the low-pressure turbine second-stage inner support ring, and the outlet end of the second pre-swirl nozzle is connected to the shared disk cavity.

[0020] Furthermore, an air intake hole is provided on the turbine disk of the first stage rotor of the low-pressure turbine. The air intake hole is used to introduce part of the airflow discharged from the outlet end of the second pre-swirl nozzle into the first stage moving blade of the low-pressure turbine. An exhaust outlet is provided at the tip of the first stage moving blade of the low-pressure turbine.

[0021] Furthermore, the second air intake pipe is connected to the inner cavity of the second stage guide vane of the low-pressure turbine through the outer ring of the low-pressure turbine. The outer ring of the low-pressure turbine is provided with an exhaust port at the position corresponding to the tip of the first stage moving blade of the low-pressure turbine, which is used to form a sealing airflow for the blade tip gap with the cooling airflow discharged from the exhaust outlet at the tip of the first stage moving blade of the low-pressure turbine.

[0022] Furthermore, the outer ring of the low-pressure turbine is also provided with an exhaust port at the position corresponding to the tip of the second-stage rotor blade of the low-pressure turbine.

[0023] To achieve the above technical effects, the present invention also provides a control method for a variable flow air system, the method being based on the aforementioned variable flow air system, comprising:

[0024] Construct a simulation model of an aero-engine that includes the aforementioned variable flow air system;

[0025] Using the incoming flow parameters and fuel supply parameters under the test conditions as input, the simulation model is used to simulate and analyze the cooling efficiency of the first stage moving blade of the low-pressure turbine under different opening degrees, as well as the engine exhaust temperature under the corresponding opening degree.

[0026] Using the cooling efficiency of the first stage moving blade of the low-pressure turbine at 100% opening of the variable cooling gas control device, the temperature ratio of the low-pressure turbine of the aero-engine, the bleed air temperature of the second bleed air pipe, the allowable temperature of the first stage moving blade of the low-pressure turbine, and the opening of the variable cooling gas control device as inputs, and the engine exhaust temperature at the corresponding opening as outputs, a relationship model between engine exhaust temperature and the opening of the variable cooling gas control device is constructed.

[0027] Based on the exhaust temperature control value of the aero-engine under the test conditions, the opening degree of the variable cooling gas control device that meets the corresponding exhaust temperature control value is obtained by using the relationship model analysis, so as to adjust the opening degree of the variable cooling gas control device.

[0028] Furthermore, the relational model is as follows: ,in The opening degree of the variable air conditioning control device, This refers to the cooling efficiency of the first-stage moving blades of the low-pressure turbine at 100% opening of the variable cooling gas control device. It is a natural constant. This is the flow coefficient correction value for the variable air conditioning control device. This is a correction factor for the blade cooling method. , Obtained through data fitting; The exhaust temperature of an aircraft engine. For the low-pressure turbine temperature ratio of aero engines, The bleed air temperature of the second bleed air tube. This refers to the allowable temperature of the first-stage moving blades of the low-pressure turbine.

[0029] Compared with the prior art, the beneficial effects of the present invention are: the present invention can continue to cool the inner support ring and other components when the cooling airflow changes and the pressure in the disk cavity decreases, so as to avoid the components from overheating and failing, maintain the cooling function of the air system components, and solve the problems of blade gas backflow, large changes in rotor axial force, and overheating of wheel rim components during the variable flow process. Attached Figure Description

[0030] Figure 1 This is a schematic diagram of the variable flow air system in the embodiment;

[0031] Figure 2 This is a schematic diagram of the airflow direction drawn out by the first air intake tube in the embodiment;

[0032] Figure 3 This is a schematic diagram of the airflow direction drawn out by the second air intake pipe in the embodiment;

[0033] The components are as follows: 1. First air intake pipe; 2. First stage guide vane of low-pressure turbine; 3. First exhaust port; 4. Inner support ring of first stage of low-pressure turbine; 5. High-pressure turbine rotor; 6. First stage rotor of low-pressure turbine; 7. First pre-swirl nozzle; 8. Second air intake pipe; 9. Second stage guide vane of low-pressure turbine; 10. Front rim jet hole; 11. Inner support ring of second stage of low-pressure turbine; 12. Rear rim jet hole; 13. Second stage rotor of low-pressure turbine; 14. Variable cooling gas control device; 15. Shared disk cavity; 16. Second pre-swirl nozzle; 17. Air intake port; 18. Outer ring of low-pressure turbine. Detailed Implementation

[0034] The present invention will now be described in further detail with reference to the embodiments and accompanying drawings. However, this should not be construed as limiting the scope of the above-described subject matter of the present invention to the following embodiments; all technologies implemented based on the content of the present invention fall within the scope of the present invention.

[0035] Example

[0036] See Figures 1 to 3A variable flow air system, comprising:

[0037] The first air intake pipe 1 has its inlet end connected to the intermediate stage of the compressor and its outlet end connected to the inner cavity of the first stage guide vane 2 of the low-pressure turbine.

[0038] The first exhaust port 3 is disposed on the inner support ring 4 of the first stage of the low-pressure turbine, and the outlet end of the first exhaust port 3 faces the upstream high-pressure turbine rotor 5. The inner support ring 4 of the first stage of the low-pressure turbine is also provided with a first pre-swirl nozzle 7 facing the downstream low-pressure turbine first-stage rotor 6. The part used to introduce the first air intake pipe 1 into the inner cavity of the guide vane 2 of the first stage of the low-pressure turbine is discharged by the cooling airflow to the upstream high-pressure turbine rotor 5, and the disk cavity between the high-pressure turbine rotor 5 and the inner support ring 4 of the first stage of the low-pressure turbine.

[0039] The second air intake pipe 8 has its inlet end connected to the intermediate stage of the compressor and its outlet end connected to the inner cavity of the second stage guide vane 9 of the low-pressure turbine.

[0040] The front rim jet hole 10 is provided on the inner support ring 11 of the second stage of the low-pressure turbine, and the outlet end of the front rim jet hole 10 faces the rim of the first stage rotor 6 of the low-pressure turbine upstream.

[0041] The rear rim jet hole 12 is provided on the inner support ring 11 of the second stage of the low-pressure turbine, and the outlet end of the rear rim jet hole 12 faces the rim of the downstream second stage rotor 13 of the low-pressure turbine.

[0042] The variable cooling air control device 14 is installed on the second air intake pipe 8 and is used to adjust the cooling air intake flow rate of the second air intake pipe 8.

[0043] In this embodiment, cooling airflow is drawn from the intermediate stage of the compressor through the first air intake pipe 1. After passing through the inner cavity of the first-stage guide vane 2 of the low-pressure turbine, the cooling airflow is discharged into the high-pressure turbine rotor 5 through the first exhaust port 3, and discharged into the main flow channel from the front of the first-stage rotor 6 of the low-pressure turbine through the first pre-swirl nozzle 7. Cooling airflow is drawn from the intermediate stage of the compressor through the second air intake pipe 8. After passing through the variable cooling air control device 14, the cooling airflow reaches the inner support ring 11 of the second stage of the low-pressure turbine through the inner cavity of the second-stage guide vane 9 of the low-pressure turbine. It is then jetted through the front rim jet hole 10 and the rear rim jet hole 12 to the space between the first-stage moving blade and the second-stage guide vane 9 of the low-pressure turbine, and then discharged into the main flow channel along the rim and the first-stage moving blade. This allows for continued cooling of components such as the inner support ring even when the cooling airflow changes or the pressure inside the disk cavity decreases, preventing components from overheating and failing, maintaining the cooling function of the air system components, and solving the problems of blade gas backflow, large changes in rotor axial force, and overheating of rim components during variable flow processes.

[0044] In this embodiment, the cooling film holes of the first-stage guide vane 2 of the low-pressure turbine are arranged on the blade basin side and the blade back cheek area. There is no cooling airflow out of the leading edge and trailing edge of the first-stage guide vane 2 of the low-pressure turbine, which ensures that the cooling airflow of the first-stage guide vane 2 of the low-pressure turbine can continue to generate circulating work downstream.

[0045] In this embodiment, a shared disk cavity 15 is also included. The shared disk cavity 15 is formed by the low-pressure turbine second-stage inner support ring 11, the low-pressure turbine first-stage rotor 6, and the low-pressure turbine second-stage rotor 13. A second pre-swirl nozzle 16 is also provided on the low-pressure turbine second-stage inner support ring 11, and the outlet end of the second pre-swirl nozzle 16 is connected to the shared disk cavity 15. The shared disk cavity 15 has grates arranged at the outlet of the second pre-swirl nozzle 16. The shared disk cavity 15 can, when the low-pressure turbine cooling airflow changes or the pressure inside the disk cavity decreases, achieve pressure balance to ensure that the axial force acting on the rotor changes very little or not at all, thus avoiding large changes in the axial force acting on the rotor and maintaining the axial force adjustment function of the air system rotor.

[0046] In this embodiment, the turbine disk of the first-stage rotor 6 of the low-pressure turbine is provided with an air intake hole 17. The air intake hole 17 is used to introduce part of the airflow discharged from the outlet end of the second pre-swirl nozzle 16 into the first-stage moving blade of the low-pressure turbine. An exhaust outlet is provided at the tip of the first-stage moving blade of the low-pressure turbine. The cooling airflow is discharged into the main flow channel from the top of the blades and other positions, which can prevent the combustion gas from entering the internal cooling channel of the blade when the cooling airflow changes or the pressure in the disk cavity decreases, thus maintaining the cooling function of the air system blades.

[0047] In this embodiment, the second exhaust pipe 8 is connected to the inner cavity of the second-stage guide vane 9 of the low-pressure turbine via the low-pressure turbine outer ring 18. The low-pressure turbine outer ring 18 has an exhaust port at the position corresponding to the blade tip of the first-stage rotor blade of the low-pressure turbine. This exhaust port, together with the cooling airflow discharged from the exhaust port at the blade tip of the first-stage rotor blade, forms a sealing airflow over the blade tip gap. On one hand, the sealing airflow cools the blade tip, reducing its temperature and extending its service life. On the other hand, it ensures the uniformity and stability of the blade tip gap and effectively reduces gas leakage at the blade tip gap, improving turbine efficiency. Similarly, in this embodiment, an exhaust port may also be provided on the low-pressure turbine outer ring 18 at the position corresponding to the blade tip of the second-stage rotor 13 of the low-pressure turbine.

[0048] Based on the same inventive concept, this embodiment also provides a control method for a variable flow air system, which, based on the aforementioned variable flow air system, includes:

[0049] S1. Construct a simulation model of an aero-engine that includes the aforementioned variable flow air system;

[0050] S2. Using the incoming flow parameters and fuel supply parameters under the test conditions as input, the simulation model is used to analyze and obtain the cooling efficiency of the first stage moving blade of the low-pressure turbine under different opening degrees, and the engine exhaust temperature under the corresponding opening degree.

[0051] S3. Using the cooling efficiency of the first stage moving blade of the low-pressure turbine at 100% opening of the variable cooling gas control device 14, the temperature ratio of the low-pressure turbine of the aero-engine, the bleed air temperature of the second bleed air pipe 8, the allowable temperature of the first stage moving blade of the low-pressure turbine, and the opening of the variable cooling gas control device 14 as inputs, and the engine exhaust temperature at the corresponding opening as output, a relationship model between the engine exhaust temperature and the opening of the variable cooling gas control device 14 is constructed.

[0052] S4 uses the relationship model to analyze and obtain the opening degree of the variable cooling gas control device 14 that meets the corresponding exhaust temperature control value based on the exhaust temperature control value of the aero-engine under the test conditions, so as to adjust the opening degree of the variable cooling gas control device 14.

[0053] In this embodiment, by simulating the operating state of an aero-engine under different operating conditions, key parameters such as the cooling efficiency of the first-stage moving blades of the low-pressure turbine and the engine exhaust temperature are obtained under different opening degrees of the variable air control device 14. This allows for the construction of a relationship model between engine exhaust temperature and opening degree. This model fully considers the complex influence of parameters such as the opening degree of the variable air control device 14, the low-pressure turbine temperature ratio of the aero-engine, and the bleed air temperature of the second bleed air pipe 8 on the engine exhaust temperature, accurately reflecting the law of change in engine exhaust temperature with the opening degree of the variable air control device 14. In engineering practice, the constructed relationship model can be used to quickly analyze and determine the required opening degree of the variable air control device 14, thereby achieving precise adjustment of the opening degree of the variable air control device 14. This allows for matching the most economical cooling airflow according to the engine's thermal load state, thereby reducing the bleed air volume of the engine's cruise air system, improving the overall engine cycle efficiency, and reducing engine fuel consumption.

[0054] In this embodiment, the relational model is: ,in The opening degree of the variable cooling air control device 14, The cooling efficiency of the first-stage moving blades of the low-pressure turbine at 100% opening is measured by the variable cooling gas control device 14. It is a natural constant. The flow coefficient correction value for the variable cooling air control device 14. This is a correction factor for the blade cooling method. , Obtained through data fitting; The exhaust temperature of an aircraft engine. For the low-pressure turbine temperature ratio of aero engines, The bleed temperature of the second bleed tube 8. This refers to the allowable temperature of the first-stage moving blades of the low-pressure turbine.

[0055] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A variable flow air system, characterized in that, include: The first air intake pipe has its inlet end connected to the intermediate stage of the compressor and its outlet end connected to the inner cavity of the first stage guide vane of the low-pressure turbine. The first exhaust port is provided on the inner support ring of the first stage of the low-pressure turbine, and the outlet end of the first exhaust port faces the upstream high-pressure turbine rotor. The inner support ring of the first stage of the low-pressure turbine is also provided with a first pre-swirl nozzle facing the downstream low-pressure turbine first stage rotor. The second air intake pipe has its inlet end connected to the intermediate stage of the compressor and its outlet end connected to the inner cavity of the guide vane of the second stage of the low-pressure turbine. The front rim jet hole is provided on the inner support ring of the second stage of the low-pressure turbine, and the outlet end of the front rim jet hole faces the rim of the first stage rotor of the low-pressure turbine upstream. The rear rim jet hole is provided on the inner support ring of the second stage of the low-pressure turbine, and the outlet end of the rear rim jet hole faces the rim of the downstream second stage rotor of the low-pressure turbine. A variable cooling air control device is installed on the second air intake pipe to adjust the cooling airflow rate of the second air intake pipe; It also includes a shared disk cavity, which is formed by the low-pressure turbine second-stage inner support ring, the low-pressure turbine first-stage rotor, and the low-pressure turbine second-stage rotor; a second pre-swirl nozzle is also provided on the low-pressure turbine second-stage inner support ring, and the outlet end of the second pre-swirl nozzle is connected to the shared disk cavity.

2. The variable flow air system according to claim 1, characterized in that, The cooling film vents of the first-stage guide vane of the low-pressure turbine are arranged on the blade basin side and the blade back cheek area, and there is no cooling airflow at the leading edge and trailing edge of the first-stage guide vane of the low-pressure turbine.

3. The variable flow air system according to claim 1, characterized in that, The turbine disk of the first stage rotor of the low-pressure turbine is provided with an air intake hole, which is used to introduce part of the airflow discharged from the outlet end of the second pre-swirl nozzle into the first stage moving blade of the low-pressure turbine. An exhaust outlet is provided at the tip of the first stage moving blade of the low-pressure turbine.

4. The variable flow air system according to claim 3, characterized in that, The second air intake pipe is connected to the inner cavity of the second stage guide vane of the low-pressure turbine through the outer ring of the low-pressure turbine. The outer ring of the low-pressure turbine is provided with an exhaust port at the position corresponding to the blade tip of the first stage moving blade of the low-pressure turbine, which is used to form a sealing airflow for the blade tip gap with the cooling airflow discharged from the exhaust outlet at the blade tip of the first stage moving blade of the low-pressure turbine.

5. The variable flow air system according to claim 4, characterized in that, The outer ring of the low-pressure turbine also has an exhaust port at the position corresponding to the tip of the second-stage rotor blade of the low-pressure turbine.

6. A control method for a variable flow air system, the method being based on the variable flow air system according to any one of claims 1-5, characterized in that, include: Construct a simulation model of an aero-engine that includes the aforementioned variable flow air system; Using the incoming flow parameters and fuel supply parameters under the test conditions as input, the simulation model is used to simulate and analyze the cooling efficiency of the first stage moving blade of the low-pressure turbine under different opening degrees, as well as the engine exhaust temperature under the corresponding opening degree. Using the cooling efficiency of the first stage moving blade of the low-pressure turbine at 100% opening of the variable cooling gas control device, the temperature ratio of the low-pressure turbine of the aero-engine, the bleed air temperature of the second bleed air pipe, the allowable temperature of the first stage moving blade of the low-pressure turbine, and the opening of the variable cooling gas control device as inputs, and the engine exhaust temperature at the corresponding opening as outputs, a relationship model between engine exhaust temperature and the opening of the variable cooling gas control device is constructed. Based on the exhaust temperature control value of the aero-engine under the test conditions, the opening degree of the variable cooling gas control device that meets the corresponding exhaust temperature control value is obtained by using the relationship model analysis, so as to adjust the opening degree of the variable cooling gas control device.

7. The control method according to claim 6, characterized in that, The relational model is as follows: ,in The opening degree of the variable air conditioning control device, This refers to the cooling efficiency of the first-stage moving blades of the low-pressure turbine at 100% opening of the variable cooling gas control device. It is a natural constant. This is the flow coefficient correction value for the variable air conditioning control device. This is a correction factor for the blade cooling method. , Obtained through data fitting; The exhaust temperature of an aircraft engine. For the low-pressure turbine temperature ratio of aero engines, The bleed air temperature of the second bleed air tube. This refers to the allowable temperature of the first-stage moving blades of the low-pressure turbine.

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