A method and configuration for designing a low infrared signature containment-divert exhaust system

Abstract

This invention belongs to the technical field of low-infrared-feature bypass split-flow exhaust system design for turbofan engines, and particularly relates to a design method and configuration for a low-infrared-feature bypass split-flow exhaust system. In this method, infrared stealth is considered during the mission requirements phase of the turbofan engine design, enabling the creation of an engine with inherent infrared stealth capabilities and greater potential for infrared stealth. The bypass split-flow exhaust system designed using this method includes a fully shielded guide vane with film cooling and a central cone structure, an axisymmetric lobed convergent nozzle, and a convergent-divergent nozzle structure. The split-flow second duct structure divides the bypass airflow into two streams at the rear casing exit section. The second duct airflow mixes with the inner airflow at the exhaust system mixer exit section. After passing through the third duct structure, the third duct airflow mixes with the mainstream at the axisymmetric lobed convergent nozzle, suppressing the infrared radiation characteristics of the nozzle tail jet.

Description

Technical Field

[0001] This invention belongs to the technical field of bypass split exhaust system design for turbofan engines with low infrared signature, and particularly relates to a design method and configuration of a bypass split exhaust system with low infrared signature. Background Technology

[0002] Infrared stealth is a crucial technological goal for fourth-generation and future fighter jets, with the infrared stealth of engine exhaust systems playing a vital role. The DC turbofan engines commonly used in fighter jets are important power plants and also the primary sources of mid-wave (3-5μm) infrared radiation. With the rapid development of infrared detection and guidance technologies, future fighter jets will face searches and detection from enemy air, space, and ground-based infrared detection systems. The all-around infrared stealth performance of aircraft is therefore of paramount importance. Thus, researching the infrared stealth characteristics of fighter jet exhaust systems in the rear hemisphere is a crucial means to improve the survivability of future fighter jets.

[0003] The infrared radiation of an engine exhaust system is mainly concentrated in the rear hemisphere, including the infrared radiation from the exhaust cavity and the tail jet. Typically, the infrared stealth characteristics of an exhaust system are evaluated based on the maximum integrated infrared radiation intensity and intensity distribution on a typical plane in the rear hemisphere. Numerous scholars have conducted extensive experimental research and theoretical numerical calculations on the infrared radiation characteristics of axisymmetric exhaust systems, direct-diffusion exhaust systems, and axisymmetric and direct-diffusion exhaust systems, based on the structural symmetry of the exhaust system, obtaining the infrared radiation intensity distribution of a direct-diffusion exhaust system in a typical detection plane. With the development of space-based infrared detection technology, the infrared radiation intensity distribution and average infrared radiation intensity in the rear hemisphere of the engine also need to be considered. Constructing a hemisphere centered on the engine nozzle, the mid-wave infrared radiation source of the engine exhaust system is mainly concentrated in this rear hemisphere. Infrared radiation in the direction directly behind the nozzle is primarily radiated by the solid walls of the exhaust system cavity, while infrared radiation perpendicular to the nozzle direction is primarily radiated by the exhaust jet.

[0004] Commonly used infrared suppression measures can be mainly divided into three categories: high-temperature component shielding technology (such as irregularly shaped nozzle structures), high-temperature component wall cooling technology (such as support plates, center cones, and nozzle wall cooling), and the use of low-emissivity materials. All three infrared suppression measures can effectively reduce the infrared radiation intensity of solid components within the exhaust system cavity in the small-angle direction of the positive tail, but they have almost no suppression effect on the infrared radiation of the exhaust system's tail jet, making it difficult to improve the suppression effect on the average infrared radiation intensity of the aircraft in the rear hemisphere.

[0005] This invention discloses a design method and configuration for a bypass split exhaust system with low infrared signature. This exhaust system structure can effectively suppress the infrared radiation intensity in the rear hemisphere direction of the exhaust system, especially the radiation from the tail jet. Summary of the Invention

[0006] To address the aforementioned deficiencies in existing technologies, this invention proposes an accurate research scheme applicable to multiple infrared suppression methods: A design method for a bypass split exhaust system with low infrared signature includes the following steps: Step (1): Based on the flight mission requirements (including infrared stealth requirements), perform 0-dimensional engine constraint analysis and mission analysis in sequence; Step (2): Select the corresponding design point design parameters; Step (3): Initially provide the relevant design point parameters for the component design; Step (4): Use the engine 0D performance calculation program to calculate whether the engine thrust at the design point meets the design target requirements. If it does not meet the requirements, return to step (3) to reselect the component design point parameters. If the design target requirements are met, calculate the key geometric area of ​​the exhaust system based on the component design point parameters and the engine 0D performance calculation program, and initially construct a 3D model of the bypass split exhaust system. Step (5): Perform numerical simulation calculation on the constructed 3D model of the bypass split exhaust system. If the exhaust system thrust does not reach the design target, readjust the total pressure recovery coefficient of the exhaust system in the engine 0D performance program and return to step (4). If the target is reached, further perform numerical simulation calculation on the infrared characteristics of the 3D model of the bypass split exhaust system with infrared suppression measures. Step (6): Determine whether the infrared characteristics of the bypass split exhaust system meet the requirements. If not, return to step (3); if they meet the requirements, complete the selection of engine component design parameters and the construction of the design point bypass split exhaust system model. Step (7): Based on the existing engine design point parameters, initially give the engine non-design point adjustment parameters, and calculate whether the engine thrust for key tasks at the non-design point meets the target requirements. If it does not meet the target, readjust the non-design point adjustment parameters; if it meets the target, output the corresponding non-design point adjustment parameters, and calculate the throat of the bypass split exhaust system based on the non-design point adjustment parameters. A 8 and nozzle exit A 9. The area is then adjusted based on the 3D model of the bypass diversion exhaust system obtained in step (6). A 8 and A 9. Construct a 3D model of the non-design point bypass exhaust system; Step (8) Perform numerical simulation on the aerodynamic and infrared characteristics of the non-design point bypass diversion exhaust system. If the infrared characteristics do not meet the requirements, return to step (3) and re-define the design parameters of the component design point; if they meet the requirements, the construction of the non-design point bypass diversion exhaust system is completed. Step (9) verifies whether the thrust and total fuel consumption of all tasks of the turbofan engine meet the requirements. If not, return to step (1) to re-perform the constraint analysis of the engine; if they meet the requirements, continue to the next step. Step (10): Determine the engine size and installation characteristics analysis, and judge again whether the thrust and total fuel consumption of all tasks and the infrared characteristics of key tasks meet the requirements. If not, return to step (1) to re-perform the constraint analysis; if the assessment target requirements are met, the iterative design of the turbofan engine is completed.

[0007] As a further improvement to this technology, the design parameters of the selected design point components in step (2) include engine fan pressure ratio, compressor pressure ratio, bypass ratio, turbine inlet temperature and bypass split ratio.

[0008] As a further improvement to this technology, it includes fully shielded guide plates with film cooling, a central cone structure with film cooling, an expanding mixer, a heat shield, a split-flow second duct, a three-duct structure, an axisymmetric wave-converging nozzle, and a contracting-diverging nozzle. Multiple fully shielded guide plates are circumferentially and evenly installed between the central cone structure and the expanding mixer. The expanding mixer is enclosed by the three-duct structure, and the split-flow second duct is nested within the three-duct structure, with its front end located behind the expanding mixer. The heat shield is nested inside the split-type second duct; the axisymmetric lobe convergent nozzle is installed at the end of the split-type second duct, and the convergent-divergent nozzle structure is installed at the end of the third duct; the axisymmetric lobe convergent nozzle is located inside the convergent-divergent nozzle; the outer duct airflow passes through the split-type second duct and is divided into an inner second stream and an outer third stream. The inner duct airflow and the second stream are mixed after the expansion mixer to form the main combustion flow. The third stream passes through the third duct and then mixes with the main combustion flow at the throat section of the exhaust system through the axisymmetric lobe convergent nozzle.

[0009] As a further improvement to this technology, the expansion mixer expands along the main gas flow direction at the A5-A6 section to form an expansion channel with an expansion cone angle of 1.2°-1.9°.

[0010] As a further improvement to this technology, the ratio of the inlet area of ​​the outer duct to the inlet area of ​​the inner duct is in the range of 0.76-1.14, the ratio of the airflow of the outer duct to the airflow of the inner duct is in the range of 1.4-2.1, and the ratio of the flow rate of the third duct to the flow rate of the second duct is in the range of 4.8-7.2.

[0011] As a further improvement to this technology, air film cooling structures are arranged on both the central cone and the fully shielded guide plate structure. The low-temperature airflow from the outer bypass duct flows through the cavity of the fully shielded guide plate to the central cone. During this process, the air film cooling structures arranged on the walls of the fully shielded guide plate and the central cone cool the outer walls of the plate and the central cone. Finally, the airflow flows out through the air film cooling structure on the wall of the central cone and mixes with the main combustion flow.

[0012] As a further improvement to this technology, the axisymmetric lobe convergent nozzle should be designed with 16-20 pairs of lobes. Each pair of lobes includes a crest and a trough. The height of the crest is 0.128 to 0.192 times the nozzle exit radius, and the depth of the trough is 0.153 to 0.229 times the nozzle exit radius.

[0013] As a further improvement to this technology, the infrared characteristic calculation of the 3D exhaust system in step (6) does not change the calculation results of the original flow field numerical model. It only artificially reduces the wall temperature of the central cone and the fully shielded guide plate with the gas film cooling structure, so that the wall cooling effect of the gas film cooling structure reaches 0.5-0.7, in order to simulate the wall temperature of the central cone and the fully shielded guide plate cooled by the low temperature airflow of the bypass. The wavelet wall parameters of the axisymmetric wavelet convergent nozzle are only adjusted to 0.15 for the emissivity of the inner and outer surfaces.

[0014] As a further improvement to this technology, the non-design point adjustment parameter in step (7) is the turbine inlet temperature.

[0015] As a further improvement to this technology, the infrared stealth requirements considered in step (1) during the 0-dimensional engine design stage are mainly the requirements for infrared radiation characteristics of the engine's tail in the 0°-90° direction.

[0016] Beneficial effects:

[0017] This invention proposes a design method and configuration for a bypass split exhaust system with low infrared signature. The beneficial effects are as follows: (1) Compared with traditional engine design methods, the design method of the bypass split exhaust system of turbofan engine with low infrared signature takes infrared stealth into consideration at the mission requirement stage, which is conducive to the overall performance design of low infrared signature aircraft engine. The designed engine has natural infrared stealth capability and has greater infrared stealth potential.

[0018] (2) Compared with traditional turbofan engines and traditional direct-flow exhaust systems that can only significantly suppress infrared radiation intensity in the positive tail direction, the bypass split exhaust system designed in this invention can cool the high-temperature gas flow through a large amount of cold air flow in the third bypass under the condition that the thrust meets the standard, thereby achieving suppression of infrared radiation intensity in all directions of the rear hemisphere. Attached Figure Description

[0019] Figure 1 This is a design process diagram; Figure 2 This is the overall design drawing of the bypass split-flow exhaust system of a turbofan engine; Figure 3 This is a diagram of the components of the bypass exhaust system; Figure 4 This is a schematic diagram of a baseline axisymmetric exhaust system; Figure 5 It refers to the assembly relationship between the split-type second duct body and the heat insulation screen in the outer duct diversion exhaust system; Figure 6 (a), (b), and (c) are the lobe structure diagrams of the axisymmetric lobe convergent nozzle of the bypass split exhaust system; Figure 7 (a), (b), and (c) are the front view, left view, and top view of the air film cooling structure on the wall of the fully shielded guide plate in the outer bypass exhaust system, respectively; (d) is a schematic diagram of the included angle of the air film holes. Figure 8 (a) and (b) are the front view and left view of the film cooling structure on the central cone wall of the bypass exhaust system, respectively. Figure 9 This is a schematic diagram of the infrared radiation characteristic distribution of the bypass split exhaust system and the reference axisymmetric exhaust system.

[0020] Label Names: 1. Outer duct; 2. Inner duct; 3. Expanding mixer; 4. Second duct body with split flow; 5. Three-duct body; 6. Heat shield; 7. Axisymmetric lobed converging nozzle; 8. Converging-diverging nozzle; 9. Fully shielded guide plate; 10. Central cone; 11. Lobe crest; 12. Lobe trough; 13. Film film orifice; 14. Straight support plate; 15. Afterburner; 16. Crest height; 17. Trough depth. Specific implementation methods

[0021] The specific embodiments of the present invention will be described in further detail below with reference to the accompanying drawings and examples. The following embodiments and drawings are used to illustrate the present invention, but are not intended to limit the scope of protection of the present invention.

[0022] as follows Figure 1 A design method for a bypass split exhaust system for a turbofan engine with low infrared signature includes the following steps: Step (1): Based on the flight mission requirements (including infrared stealth requirements), perform 0-dimensional engine constraint analysis and mission analysis in sequence to determine the overall engine performance under each flight mission and preliminarily determine the adjustment range of each engine parameter. Step (2) Select the corresponding design point design parameters, which are applicable to the final design of the engine; Step (3): Initially provide the relevant design point parameters for the component design, and establish a 0-dimensional engine model as the starting engine for iterative engine design; Step (4): Using the engine 0D performance calculation program described in the reference document "Modeling of the turbofan with an ejector nozzle based on infrared prediction", calculate whether the engine thrust at the design point meets the design target requirements. If it does not meet the requirements, return to step (3) to reselect the component design point parameters. If the design target requirements are met, calculate the key geometric area of ​​the exhaust system based on the component design point parameters and the engine 0D performance calculation program, and initially construct a 3D model of the bypass split exhaust system. Step (5): Perform numerical simulation calculation on the constructed 3D model of the bypass split exhaust system. If the exhaust system thrust does not reach the design target, readjust the total pressure recovery coefficient of the exhaust system in the engine 0D performance program and return to step (4). If the target is reached, further perform numerical simulation calculation on the infrared characteristics of the 3D model of the bypass split exhaust system with infrared suppression measures. Step (6) determines whether the infrared characteristics of the bypass split exhaust system meet the requirements. If not, return to step (3); if they meet the requirements, the selection of engine component design parameters and the construction of the bypass split exhaust system model at the design point are completed. Steps (4) to (6) complete the iterative process of the overall performance design of the engine and the numerical simulation of the aerodynamic and infrared characteristics of the three-dimensional exhaust system at the design point. Step (7): Based on the existing engine design point parameters, initially give the engine non-design point adjustment parameters, and calculate whether the engine thrust for key tasks at the non-design point meets the target requirements. If it does not meet the target, readjust the non-design point adjustment parameters; if it meets the target, output the corresponding non-design point adjustment parameters, and calculate the throat of the bypass split exhaust system based on the non-design point adjustment parameters. A 8 and nozzle exit A 9. The area is then adjusted based on the 3D model of the bypass diversion exhaust system obtained in step (6). A 8 and A 9. Construct a 3D model of the non-design point bypass exhaust system; Step (8) performs numerical simulation of the aerodynamic and infrared characteristics of the bypass split exhaust system at non-design points. If the infrared characteristics do not meet the requirements, return to step (3) and re-define the design parameters of the component design point. If the requirements are met, the construction of the bypass split exhaust system at non-design points is completed. Step (8) completes the iterative process of the overall engine performance design and the numerical simulation of the aerodynamic and infrared characteristics of the three-dimensional exhaust system in the non-design point state. Step (9) verifies whether the thrust and total fuel consumption of all tasks of the turbofan engine meet the requirements. If not, return to step (1) to re-perform the constraint analysis of the engine; if they meet the requirements, continue to the next step. Step (10): Determine the engine size and installation characteristics analysis, and judge again whether the thrust and total fuel consumption of all tasks and the infrared characteristics of key tasks meet the requirements. If not, return to step (1) to re-perform the constraint analysis; if the assessment target requirements are met, the iterative design of the turbofan engine is completed.

[0023] Figure 2 The diagram shows the overall design of the bypass split-flow exhaust system for a turbofan engine. Section A5 is the inlet section for the inner airflow, section A6 is the outlet section for the inner airflow mixer 3, section A15 is the inlet section for the outer airflow, section A16 is the second stream split-flow section, section A26 is the third stream split-flow section, section A7 is the inlet section for the main combustion gas flow in the nozzle, section A8 is the throat section for the main combustion gas flow, and section A9 is the outlet section for the nozzle. m 6 represents the internal airflow. m 16 and m 26 The sum of these is the external bypass airflow, where m 16 For the second duct airflow, m 26 The flow rate of the third duct is the flow rate of the outer duct. m 26 With the second duct flow m 16 ratio.

[0024] Figure 3 The diagram shows the components of the ducted exhaust system. The inner airflow, which is high-temperature air, flows into the exhaust system from the inner channel 2. The fully shielded guide plate 9 is a torsion plate that, through a reasonable surface design, completely shields the high-temperature components in front of section A5 from the exhaust system's tail direction. An expansion mixer 3 is installed on the outer side of the fully shielded guide plate 9. A central cone 10 is installed on the inner side of the fully shielded guide plate 9. The wall of the central cone 10 and the inner wall of the expansion mixer 3 form the inner channel 2. The outer airflow flows into the exhaust system from the outer channel 1. The outer airflow is divided into the second duct airflow in the split-type second duct body 4. m 16 and the third duct airflow m 26 Internal airflow m 6. Second duct airflow m 16 The main gas flow is formed by mixing after the expansion mixer 3; the third duct gas flow... m 26The gas flow is mixed with the main gas flow through the axisymmetric lobe convergent nozzle 7; Figure 4 The figure shows a reference axisymmetric exhaust system, including a central cone 10, a common straight support plate 14, an afterburner 15, and a converging nozzle 8. It has an inner airflow 2 and an outer airflow 1. The two airflows are mixed in the afterburner 15 and finally discharged from the converging nozzle 8.

[0025] Figure 5 (a) shows the expansion angle of the expansion mixer 3 of the bypass split exhaust system, which is designed to be 1.57°. This angle design can greatly reduce the flow resistance of the bypass airflow. Figure 5 (b) shows the assembly relationship between the second duct body 4 and the heat shield 6 of the duct diversion exhaust system. The second duct body 4 wraps around the heat shield 6, and the heat shield 6 extends all the way to the tail of the second duct body 4.

[0026] Figure 6 (a) shows a schematic diagram of the lobe structure of the axisymmetric lobe convergent nozzle 7 in the bypass split exhaust system, with a larger flow rate in the third bypass airflow. m 26 After passing through the three-duct cylinder 5, the main gas flow is mixed with the axisymmetric lobe convergent nozzle 7, which suppresses the infrared radiation characteristics of the nozzle tail jet. Figure 6 (b) and (c) are the front view and side view of the specific structure, with a total of 16 pairs of identical lobe structures. The crest 11 and trough 12 of the lobe are both rectangular. The included angle between the walls of each pair of crest 11 and trough 12 structures is 11.25°. The height 16 of the crest is 0.16 times the nozzle exit radius, and the depth 17 of the trough is 0.19 times the nozzle exit radius.

[0027] Figure 7 Figures (a), (b), and (c) show the structure of the film cooling system on the wall of the fully shielded guide plate 9. The film cooling holes 13 are arranged in a staggered pattern, and the diameter of the film cooling holes 13 is 0.4 mm. Figure 7 (d) shows that the angle between the air film hole 13 and the wall of the fully shielded flow guide plate 9 is 20°.

[0028] Figure 8 The figure shown is a diagram of the film cooling structure 13 on the wall of the central cone 10. The diameter of the film cooling hole 13 is 0.4 mm, and the corresponding hole inclination angle is 90° perpendicular to the wall of the central cone 10. The film cooling holes 13 of the central cone 10 are arranged in a staggered manner.

[0029] The following numerical simulation example illustrates the effect of suppressing the rear hemisphere infrared radiation of an exhaust system designed using the bypass split exhaust system design method with low infrared characteristics disclosed in this invention, compared to a reference axisymmetric exhaust system with the same engine thrust.

[0030] like Figure 9The figure shows the infrared integrated radiation intensity distribution of the reference axisymmetric exhaust system and the bypass split-flow exhaust system within the horizontal detection plane in the 3~5μm band during supercruise. The solid black line represents the reference axisymmetric exhaust system, the dashed line represents the infrared radiation intensity distribution of the bypass split-flow exhaust system without other infrared suppression measures, and the dotted line represents the infrared radiation intensity distribution of the bypass split-flow exhaust system after the use of a fully shielded guide plate 9 and a central cone 10 achieving 60% wall cooling, and the application of a low emissivity coating of 0.15 to the lobe structure of the axisymmetric lobe convergent nozzle 7. Without suppression measures, the infrared radiation intensity of the bypass split-flow exhaust system at 0° decreased by 4.25% relative to the reference axisymmetric system. This is because the lobe structure of the axisymmetric lobe convergent nozzle 7 extends deep into the core gas flow, blocking part of the solid radiation of the exhaust system at 0°. Furthermore, the lobe of the axisymmetric lobe convergent nozzle 7 is cooled by the third bypass airflow, resulting in a lower wall temperature. Overall, there was no significant reduction in the suppression effect on the infrared radiation intensity at 0°. The infrared radiation intensity at 90° decreased by 69.37% relative to the reference axisymmetric exhaust system. This is due to the three-channel airflow... m 26 The mixing effect with mainstream gas is significantly enhanced, greatly reducing the jet temperature and significantly decreasing the jet infrared radiation. The average infrared radiation intensity in the rear hemisphere decreases by 61.55% compared to the reference axisymmetric exhaust system. By employing wall cooling measures on the fully shielded guide plate 9 and the central cone 10, and by additionally applying a low-emissivity coating η=0.15 to the lobe structure of the axisymmetric lobe convergent nozzle 7, the infrared radiation intensity of the exhaust system decreases significantly within the 0~15° range. The maximum infrared radiation intensity at 0° decreases by 72.3% compared to the reference axisymmetric exhaust system; the infrared radiation intensity at 90° remains unchanged, but the rear hemisphere decreases by 71.9% relative to the reference. These examples demonstrate that the design method and exhaust system of this invention effectively suppress the infrared radiation intensity in the rear hemisphere.

Claims

1. A design method for a bypass split exhaust system with low infrared signature, characterized in that, Includes the following steps: Step (1): Based on the flight mission requirements (including infrared stealth requirements), perform 0-dimensional engine constraint analysis and mission analysis in sequence; Step (2): Select the corresponding design point design parameters; Step (3): Initially provide the relevant design point parameters for the component design; Step (4): Use the engine 0D performance calculation program to calculate whether the engine thrust at the design point meets the design target requirements. If it does not meet the requirements, return to step (3) to reselect the component design point parameters. If the design target requirements are met, calculate the key geometric area of ​​the exhaust system based on the component design point parameters and the engine 0D performance calculation program, and initially construct a 3D model of the bypass split exhaust system. Step (5): Perform numerical simulation calculation on the constructed 3D model of the bypass split exhaust system. If the exhaust system thrust does not reach the design target, readjust the total pressure recovery coefficient of the exhaust system in the engine 0D performance program and return to step (4). If the target is reached, further perform numerical simulation calculation on the infrared characteristics of the 3D model of the bypass split exhaust system with infrared suppression measures. Step (6): Determine whether the infrared characteristics of the bypass split exhaust system meet the requirements. If not, return to step (3); if they meet the requirements, complete the selection of engine component design parameters and the construction of the design point bypass split exhaust system model. Step (7): Based on the existing engine design point parameters, initially give the engine non-design point adjustment parameters, and calculate whether the engine thrust for key tasks at the non-design point meets the target requirements. If it does not meet the target, readjust the non-design point adjustment parameters; if it meets the target, output the corresponding non-design point adjustment parameters, and calculate the throat of the bypass split exhaust system based on the non-design point adjustment parameters. A 8 and nozzle exit A 9. Area, then based on the design point bypass diversion exhaust system 3D model obtained in step (6), by adjusting A 8 and A 9. Construct a 3D model of the non-design point bypass exhaust system; Step (8) Perform numerical simulation on the aerodynamic and infrared characteristics of the non-design point bypass diversion exhaust system. If the infrared characteristics do not meet the requirements, return to step (3) and re-define the design parameters of the component design point; if they meet the requirements, the construction of the non-design point bypass diversion exhaust system is completed. Step (9) verifies whether the thrust and total fuel consumption of all tasks of the turbofan engine meet the requirements. If not, return to step (1) to re-perform the constraint analysis of the engine; if they meet the requirements, continue to the next step. Step (10): Determine the engine size and installation characteristics analysis, and judge again whether the thrust and total fuel consumption of all tasks and the infrared characteristics of key tasks meet the requirements. If not, return to step (1) to re-perform the constraint analysis; if the assessment target requirements are met, the iterative design of the turbofan engine is completed.

2. The design method for a low infrared signature bypass diversion exhaust system as described in claim 1, characterized in that: The design parameters of the components selected in step (2) include engine fan pressure ratio, compressor pressure ratio, bypass ratio, turbine inlet temperature and bypass split ratio.

3. An exhaust system configuration constructed according to the low infrared signature bypass split exhaust system design method as described in claim 1, characterized in that: It includes fully shielded guide plates with film cooling, a central cone structure with film cooling, an expanding mixer, a heat shield, a split-flow second duct, a three-duct structure, an axisymmetric lobed converging nozzle, and a contracting-diverging nozzle. Multiple fully shielded guide plates are circumferentially and evenly installed between the central cone structure and the expanding mixer. The expanding mixer is enclosed by the three-duct structure. The split-flow second duct is nested within the three-duct structure, with its front end located behind the expanding mixer. The heat shield is nested within... Inside the split-flow second duct; an axisymmetric lobe convergent nozzle is installed at the end of the split-flow second duct, and a convergent-divergent nozzle structure is installed at the end of the third duct; the axisymmetric lobe convergent nozzle is located inside the convergent-divergent nozzle; the outer duct airflow passes through the split-flow second duct and is divided into an inner second stream and an outer third stream. The inner duct airflow and the second stream are mixed after the expansion mixer to form the main combustion flow. The third stream passes through the third duct and then mixes with the main combustion flow at the throat section of the exhaust system through the axisymmetric lobe convergent nozzle.

4. The exhaust system configuration as described in claim 3, characterized in that: The expansion mixer expands along the main gas flow direction at section A5-A6 to form an expansion channel with an expansion cone angle of 1.2°-1.9°.

5. An exhaust system configuration as claimed in claim 3, characterized in that: The ratio of the inlet area of ​​the outer duct to the inlet area of ​​the inner duct is in the range of 0.76-1.14, the ratio of the airflow rate of the outer duct to the airflow rate of the inner duct is in the range of 1.4-2.1, and the ratio of the flow rate of the third duct to the flow rate of the second duct is in the range of 4.8-7.

2.

6. The exhaust system configuration as described in claim 3, characterized in that: Both the central cone and the fully shielded guide plate structure are equipped with film cooling structures. The low-temperature airflow from the outer bypass duct flows through the cavity of the fully shielded guide plate to the central cone. During this process, the film cooling structures arranged on the walls of the fully shielded guide plate and the central cone cool the outer walls of the plate and the central cone. Finally, the airflow flows out through the film cooling structure on the wall of the central cone and mixes with the main combustion flow.

7. The exhaust system configuration as described in claim 3, characterized in that: The axisymmetric convergent nozzle should be designed with 16-20 pairs of lobes. Each pair of lobes consists of a crest and a trough. The crest height is 0.128 to 0.192 times the nozzle exit radius, and the trough depth is 0.153 to 0.229 times the nozzle exit radius.

8. The design method for a low infrared signature bypass diversion exhaust system as described in claim 1, characterized in that: In step (6), the infrared characteristics calculation of the 3D exhaust system does not change the calculation results of the original flow field numerical model. Only the wall temperature of the central cone and the fully shielded guide plate with the gas film cooling structure is artificially reduced so that the wall cooling effect of the gas film cooling structure reaches 0.5-0.7, in order to simulate the wall temperature of the central cone and the fully shielded guide plate cooled by the low temperature airflow of the bypass. The wavelet wall parameters of the axisymmetric wavelet convergent nozzle are adjusted only by adjusting the emissivity of the inner and outer surfaces to 0.

15.

9. The design method for a low infrared signature bypass diversion exhaust system as described in claim 1, characterized in that: In step (7), the non-design point adjustment parameter is the turbine inlet temperature.

10. The design method for a low infrared signature bypass split exhaust system as described in claim 1, characterized in that: In step (1), the infrared stealth requirements considered in the 0-dimensional engine design stage are mainly the infrared radiation characteristics requirements of the engine's tail in the 0° to 90° direction.

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