Helicopter infrared suppressor fairing and gas heat optimization design method thereof

Abstract

The invention belongs to the technical field of overall aerodynamic design of helicopters, and particularly relates to an infrared suppressor fairing of a helicopter and an air heat optimization design method of the infrared suppressor fairing. The fairing comprises a fairing body, a shutter air inlet, a square protective net air inlet, a second-stage exhaust pipe and a first-stage mixed exhaust pipe. The front of the fairing body is regarded as a half ellipse which is laterally cut along a long axis; a shutter air inlet is formed in the front portion of the fairing body, and the plane of a shutter is crescent A square protective net air inlet is formed in front of the top of the fairing body; an oval second-stage exhaust pipe is arranged behind the top of the fairing body so that high-temperature exhaust can be further cooled by means of downwash airflow of the rotor wings. A first-stage mixed exhaust pipe is sleeved in the second-stage exhaust pipe; and the primary mixing exhaust pipe is connected with an outlet of the transmitter. According to the method, the problem that the pneumatic performance and the thermal performance are difficult to consider in a traditional design method is solved, comprehensive optimization of the pneumatic performance and the thermal performance of the fairing is achieved, and a pneumatic design scheme of the infrared suppressor fairing is formed.

Description

Technical Field

[0001] The invention belongs to the technical field of overall aerodynamic design of helicopters, and in particular relates to a helicopter infrared suppressor fairing and an aero-thermal optimization design method thereof. Background Art

[0002] Infrared suppressors are key devices for reducing the infrared target radiation signature of helicopters and improving their stealth capabilities. As a crucial component of an infrared suppressor, the fairing's aerodynamic shape and thermal characteristics directly impact its overall performance. Traditional fairing design methods often separate aerodynamic and thermal design, making it difficult to achieve comprehensive optimization of aerodynamic and thermal performance. They often focus on structural strength and thermal protection, while insufficiently optimizing aerodynamic performance. This results in significant aerodynamic drag during flight, or the fairing's shedding vortex adversely interfering with the tail surface, impacting the helicopter's overall performance and limiting the use of infrared suppressors. Summary of the Invention

[0003] Purpose of the invention: The present invention aims to propose a helicopter infrared suppressor fairing and its aero-thermal optimization design method, to solve the problem that it is difficult to take both aerodynamic and thermal performance into consideration in traditional design methods, to achieve comprehensive optimization of the fairing's aerodynamic and thermal performance, and to form an aerodynamic design scheme for the infrared suppressor fairing.

[0004] Technical solution:

[0005] In a first aspect, a helicopter infrared suppressor fairing is provided, comprising: a fairing body, a shutter air inlet, a square protective net air inlet, a secondary exhaust pipe, and a primary mixing exhaust pipe;

[0006] The fairing body is viewed from the front as a half ellipse cut along the long axis; the half ellipse is composed of two different upper 1 / 4 ellipse and lower 1 / 4 ellipse;

[0007] The front of the fairing body is provided with a louver air inlet, the louver plane is crescent-shaped; the angle between the louver plane and the longitudinal symmetry plane is 35°~40°; the louver blades are L-shaped, the angle between the outer edge of the blade and the horizontal plane is 18°~20°, and the angle between the inner edge of the blade and the horizontal plane is 45°~47°;

[0008] A square protective mesh air inlet is provided at the front of the top of the fairing body;

[0009] An elliptical secondary exhaust pipe is located at the rear of the top of the fairing body to further cool the high-temperature exhaust gas with the help of the rotor downwash airflow. The primary mixing exhaust pipe is housed within the secondary exhaust pipe, which is connected to the engine outlet. The angle between the secondary exhaust pipe axis and the horizontal plane is 45° to 50°. The gap between the secondary exhaust pipe and the primary mixing exhaust pipe is 40 to 45 mm, and the overlap between the secondary exhaust pipe and the primary mixing exhaust pipe is 50 mm to 150 mm.

[0010] The rear of the fairing is designed with a streamlined tail cone. The shape of the streamlined tail cone is composed of two parabolas, and the side gradually transitions from an ellipse to a hyperbolic shape. There is a gap between the tail cone and the fuselage skin to facilitate the exhaust of gas inside the infrared suppressor fairing cabin and the intake of external cold air, thereby reducing the cabin temperature. The tail cone aspect ratio γ is 1 to 1.5.

[0011] The air flow enters from the louver air inlet and the square protective net air inlet, and is discharged from the gap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe through the high-speed airflow injection effect of the first-stage mixing exhaust pipe; the fairing body does not directly contact the first-stage mixing exhaust pipe, and the distance between the fairing body and the first-stage mixing exhaust pipe is maintained at 150mm~200m.

[0012] In a second aspect, a method for aero-thermal optimization design of a helicopter infrared suppressor fairing is provided, comprising:

[0013] The infrared suppressor fairing adopts a convex leading edge design to increase the windward area, create a wind-driving effect, and improve the air intake efficiency;

[0014] The trailing edge of the fairing is tapered to reduce aerodynamic drag;

[0015] The side of the fairing adopts a multi-section curved surface splicing design, and the curvature of each section of the surface is optimized according to the flow field distribution;

[0016] An internal airflow channel is designed between the infrared suppressor fairing and the first-stage mixing exhaust pipe to guide the cooling airflow into the first-stage mixing exhaust pipe for mixing;

[0017] A louver air inlet and a square protective net air inlet are designed on the infrared suppressor fairing, so that air enters from the louver air inlet and the square protective net air inlet, and is discharged from the gap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe through the high-speed airflow injection effect of the first-stage mixing exhaust pipe.

[0018] Furthermore, the design method is implemented in the following manner:

[0019] Based on the dimensions of the helicopter engine exhaust port and the internal structure of the infrared suppressor, a parametric model of the infrared suppressor fairing was established. The parameters of the parametric model included the major and minor axes of the elliptical curve of the fairing's side, and the aspect ratio of the tail cone.

[0020] The purpose of model optimization is to reduce the aerodynamic drag coefficient of the fairing, weaken the aerodynamic interference of the fairing on the horizontal tail, and reduce the surface temperature of the fairing and the average temperature of the tail exhaust.

[0021] Based on the internal and external flow fields and parameterized models of CFD methods, a simulation analysis model that fully considers the fuselage and engine exhaust system is established;

[0022] The unstructured tetrahedron grid is used to perform aerodynamic simulation calculations on the simulation analysis model, and the grid is encrypted in the areas where the geometric curvature of the simulation analysis model changes greatly and the flow field changes violently to improve the accuracy of flow field capture; the momentum source method is used to simulate the rotor downwash flow field; the rotor downwash flow field is part of the external flow field; the ROE flux difference splitting format is used for the spatial discretization of the simulation calculation, and the two-equation turbulence model used in the simulation calculation is used. Realizable model, using the DO model for thermal radiation used in simulation calculations; the external flow field uses velocity inlet and pressure outlet boundary conditions, the engine exhaust uses the internal flow field flow inlet boundary, and the object surface uses a no-slip boundary;

[0023] Based on the aerodynamic simulation results, flow field velocity and temperature field distribution, the aerodynamic and thermal characteristics analysis of the infrared suppressor fairing is carried out.

[0024] Furthermore, the aerodynamic and thermal characteristics analysis of the infrared suppressor fairing was carried out, including:

[0025] By optimizing the three-dimensional shape of the fairing's sides and trailing edge, smoothing the airflow behind the fairing, reducing airflow separation, and minimizing the fairing's outer contour as much as possible, drag is reduced, weakening the adverse interference with the tail aerodynamic surface;

[0026] To improve the ejection effect, the surface temperature of the infrared suppressor fairing is reduced, and the gap and overlap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe are optimized;

[0027] The influence of the infrared suppressor fairing shape on the aerodynamic performance and thermal characteristics is comprehensively weighed to determine the aerodynamic scheme of the infrared suppressor fairing.

[0028] Beneficial effects:

[0029] The present invention provides an aero-thermal optimization design method for a helicopter infrared suppressor fairing, which can be used to comprehensively optimize the aerodynamic and thermal performance of the infrared suppressor fairing. By optimizing the shape and internal channels of the fairing, the aerodynamic drag of the entire aircraft is reduced by 5% after the fairing is optimized, thereby reducing the helicopter's cruising fuel consumption and increasing its range. The aerodynamic efficiency of the horizontal tail is increased by 30%, which is conducive to reducing the nose-down angle of the helicopter's fuselage during high-speed leveling. The surface temperature of the infrared suppressor fairing is reduced by 15°C, thereby improving the helicopter's infrared stealth performance. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] Figure 1 This is a side view of the infrared suppressor fairing involved in the present invention.

[0031] Figure 2 This is a front view of the infrared suppressor fairing involved in the present invention.

[0032] Figure 3 This is a top view of the infrared suppressor fairing involved in the present invention.

[0033] Figure 4 This is a schematic diagram of the interference of the infrared suppressor fairing on the tail surface involved in the present invention.

[0034] Figure 5 This is a schematic diagram of the temperature field inside and outside the infrared suppressor fairing involved in the present invention. DETAILED DESCRIPTION

[0035] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the implementation of this application will be described in more detail below in conjunction with the drawings in the implementation of this application. In the drawings, the same or similar numbers throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and should not be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in conjunction with the drawings.

[0036] In the description of the present invention, it should be understood that the terms "center", "axial", "vertical", "up", "down", "upper end", "bottom end", "inside", "outside", etc., indicating the orientation or position relationship, are based on the orientation or position relationship shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the scope of protection of the present invention.

[0037] The present invention proposes an aero-thermal optimization design method for a helicopter infrared suppressor fairing, which can realize the control of the flow inside and outside the infrared suppressor fairing, reduce aerodynamic drag, weaken the aerodynamic interference on the tail aerodynamic surface, and meet the requirements of the infrared suppressor induced mixing of ambient cold air to reduce the exhaust temperature, shield the high-temperature components of the exhaust system, and at the same time integrate well with the overall shape of the fuselage.

[0038] Design Principle of Infrared Suppressor Fairing

[0039] The leading edge of the infrared suppressor fairing is convex to increase the windward area, creating a wind-driving effect and improving the air intake efficiency; the trailing edge of the fairing adopts a tapered design to reduce aerodynamic resistance; the side of the fairing adopts multiple curved surfaces, and the curvature of each curved surface is optimized according to the flow field distribution to weaken the separation phenomenon caused by the airflow flowing on the fairing surface.

[0040] An internal airflow channel is formed between the infrared suppressor fairing and the first-stage mixing exhaust pipe. This channel guides cooling air into the first-stage mixing exhaust pipe for mixing, reducing exhaust temperature. Air then flows through the fairing interior, lowering the fairing surface temperature. Air enters through the louvered inlet 1 and the square protective mesh inlet 2, and is ejected through the gap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe by the high-speed airflow ejection.

[0041] Design of infrared suppressor fairing

[0042] The infrared suppressor fairing includes: a fairing body, a shutter air inlet 1, a square protective net air inlet 2, a secondary exhaust pipe 3, and a primary mixing exhaust pipe 5;

[0043] The fairing body is a semi-ellipse cut along its long axis when viewed from the front. This not only increases the gap between the first-stage mixing exhaust pipe and the fairing, reducing the heat radiation from the high-temperature exhaust pipe to the fairing, but also reduces the airflow separation of the fairing and the vertical drag of the fairing caused by rotor downwash.

[0044] The front of the fairing body is equipped with a louvered air inlet. The louver is crescent-shaped in plan. The louver can shield the high-temperature first-stage mixing exhaust pipe and other components, reducing infrared radiation characteristics. The oblique crescent-shaped air inlet can increase the wind-driving effect and reduce the intake resistance during forward flight.

[0045] A square protective mesh air inlet is provided on the top front of the fairing body, which can directly use the rotor downwash airflow velocity to fill the fairing cabin with air, increase the air intake flow in hovering or low-speed flight state, reduce the temperature in the fairing cabin, and reduce infrared radiation;

[0046] An elliptical secondary exhaust pipe is installed at the rear of the top of the fairing body to further cool the high-temperature exhaust gas with the help of the rotor downwash airflow; the secondary exhaust pipe is sheathed with a primary mixing exhaust pipe. The gap between the secondary exhaust pipe and the primary mixing exhaust pipe can be used to inject cool air into the high-temperature exhaust gas from the primary mixing exhaust pipe to reduce the exhaust temperature; the primary mixing exhaust pipe is connected to the engine outlet;

[0047] A streamlined tail cone 4 is designed at the rear of the fairing body. The streamlined tail cone is formed by splicing two parabolas, and the side surface gradually transitions from an ellipse to a hyperbolic shape. This allows the airflow above, below, and on the sides of the fairing to converge smoothly at the rear of the fairing, thereby reducing airflow separation, thereby reducing the fairing resistance and weakening the adverse interference of the fairing separation vortex on the helicopter's tail aerodynamic surface; there is a gap between the tail cone and the fuselage skin to facilitate the exhaust of gas from the infrared suppressor fairing cabin and the entry of external cold air, thereby reducing the cabin temperature;

[0048] Airflow enters through the louvered air inlet and the square protective net air inlet, and is discharged from the gap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe through the high-speed airflow injection effect. The front louvered air inlet and the upper square protective net air inlet jointly ensure that sufficient cold air enters during forward flight and hovering, which is used to mix with the high-temperature exhaust gas.

[0049] The half ellipse is composed of two different upper 1 / 4 ellipses and lower 1 / 4 ellipses. The long axis of the upper ellipse is horizontal to ensure sufficient layout space for the first-stage mixing exhaust pipe and the fairing, and the long axis of the lower ellipse is vertical to reduce airflow separation at the bottom of the fairing.

[0050] The angle between the louver plane and the longitudinal symmetry plane is 35° to 40°, which is the result of a comprehensive balance between the slanted louver plane to reduce forward flight resistance and the forward louver plane to increase air intake. The louver blades are L-shaped, with the outer edge of the blade at an angle of 18° to 20° to the horizontal plane and the inner edge at an angle of 45° to 47° to the horizontal plane. This allows the louvers to shield high-temperature components and create a swirling flow in the airflow entering the fairing cabin to enhance mixing, while ensuring that the intake resistance does not increase significantly.

[0051] The angle between the axis of the secondary exhaust pipe and the horizontal plane is 45°~50°. The upward-facing tail exhaust and the rotor downwash flow intersect vertically to more fully mix with the cold air.

[0052] The gap between the secondary exhaust pipe and the primary mixing exhaust pipe is 40-45mm, and the overlap between the secondary exhaust pipe and the primary mixing exhaust pipe is 50-150mm, which can improve the ejection capacity of the secondary exhaust pipe;

[0053] The tail cone aspect ratio γ is 1~1.5, which is the result of a comprehensive balance between the tail cone's drag reduction effect and the structural weight cost;

[0054] The fairing body is not in direct contact with the first-stage mixing exhaust pipe. The distance between the fairing body and the first-stage mixing exhaust pipe is maintained at 150mm~200m, which can reduce the heat radiation of the high-temperature first-stage mixing exhaust pipe to the fairing, while limiting the outer contour size of the fairing to reduce resistance.

[0055] like Figure 2 As shown, the infrared suppressor fairing is viewed from the front as a side-cut half ellipse, and the half ellipse is divided into an upper quarter ellipse with a major axis a1 and a minor axis b1, and a lower quarter ellipse with a major axis a2 and a minor axis b2.

[0056] like Figure 1 、 3As shown, the infrared suppressor fairing is provided with a shutter air inlet 1 along the front of the heading, and the shutter plane 31 is crescent-shaped with an angle of 35°~40° with the longitudinal symmetry plane; the shutter blades are L-shaped, with the outer edge 11 and the horizontal plane at an angle of 18°~20°, and the inner edge 12 and the horizontal plane at an angle of 45°~47°. The L-shape can pre-swirl the cooling airflow to improve the mixing efficiency, and at the same time effectively shield the high-temperature components of the exhaust system.

[0057] A square protective mesh air inlet 2 is located at the front of the infrared suppressor fairing to increase airflow during hovering. An elliptical secondary exhaust pipe 3 is located at the rear of the top, with its axis angled at 45° to 50° with the horizontal to further cool the high-temperature exhaust gas using rotor downwash. The secondary exhaust pipe 3 allows for the passage of the primary mixing exhaust pipe 5, with a clearance of 40 to 45 mm and an overlap of 50 to 150 mm. This allows the high-speed exhaust gas to be ejected from the secondary pipe, reducing the temperature within the infrared suppressor fairing.

[0058] The rear of the infrared suppressor fairing is designed with a streamlined tail cone 4, the length-to-height ratio γ of the tail cone is 1~1.5, and it is composed of two upper and lower parabolas. Figure 1 The red and green curves in the figure are stitched together, and the side profile gradually transitions from an ellipse to a hyperbolic shape. A gap exists between the tail cone and the fuselage skin to allow for the exhaust of gases from the infrared suppressor fairing and the intake of cool air, reducing the cabin temperature.

[0059] The distance between the infrared suppressor fairing and the first-stage mixing exhaust pipe 5 is maintained at 150mm~200m, and the heat radiation of the first-stage mixing exhaust pipe to the infrared suppressor fairing is reduced. In the local area where the distance between the two cannot be guaranteed due to structural limitations, insulation materials need to be configured.

[0060] Optimization design method of infrared suppressor fairing

[0061] Based on the dimensions of the helicopter engine exhaust port and the internal structure of the infrared suppressor, a parametric model of the infrared suppressor fairing was established, including geometric parameters such as the major and minor axes of the fairing's lateral elliptical curve and the aspect ratio of the tail cone. The optimization objectives were to reduce the fairing's aerodynamic drag coefficient, mitigate the fairing's aerodynamic interference with the horizontal tail, and lower the fairing's surface temperature and the average exhaust temperature.

[0062] Internal and external flow fields based on CFD method ( Figure 5 ) and parameterized models to establish a simulation analysis model that comprehensively considers the fuselage and engine exhaust system; unstructured tetrahedral grids are used for aerodynamic simulation calculations, and grid encryption is performed in areas with large changes in geometric curvature and drastic changes in the flow field to improve the accuracy of flow field capture. The total number of grids is 20 million to 25 million.

[0063] The momentum source method is used to simulate the rotor downwash flow field (part of the external flow field); the spatial discretization of the simulation calculation adopts the flux difference splitting format of ROE, and the turbulence model adopts the two-equation A realizable model was used, and the DO model was used for thermal radiation. The external flow field used velocity inlet and pressure outlet boundary conditions, the engine exhaust used the internal flow field flow inlet boundary, and the object surface used a no-slip boundary.

[0064] Based on the CFD calculation results, including flow field velocity and temperature field distribution, the aerodynamic and thermal characteristics analysis of the infrared suppressor fairing is carried out, such as Figure 4 and Figure 5 As shown in the figure, the three-dimensional shape of the fairing's sides and trailing edge was optimized, smoothing the airflow behind the fairing, reducing airflow separation, and minimizing the fairing's outer contour to reduce drag and weaken adverse interference with the tail aerodynamic surfaces. The gap and overlap between the first-stage mixing exhaust pipe 5 and the second-stage exhaust pipe 3 were optimized to improve the ejection effect and reduce the surface temperature of the infrared suppressor fairing. The aerodynamic scheme of the infrared suppressor fairing was determined by comprehensively weighing the impact of the infrared suppressor fairing's shape on aerodynamic performance and thermal characteristics.

[0065] The present invention analyzes the influence of the infrared suppressor fairing on the aerodynamic characteristics of the fuselage based on the CFD method, optimizes the aerodynamic shape of the fairing by segmented curvature to reduce aerodynamic drag, and weakens the interference of the fairing shedding vortex on the tail aerodynamic surface.

[0066] The present invention simulates the rotor downwash flow field by adopting the momentum source method and optimizes the side profile of the infrared suppressor fairing to ensure that the rotor downwash does not cause the high-temperature tail exhaust to act directly on the fairing skin, causing the outer surface temperature of the fairing to be too high and affecting the effectiveness of the infrared suppressor.

[0067] The rear part of the infrared suppressor fairing of the present invention adopts a streamlined tail cone 4 to reduce aerodynamic resistance; a gap exists between the tail cone and the fuselage skin to facilitate the discharge of gas in the infrared suppressor fairing cabin and reduce the cabin temperature.

[0068] The shape and angle of the guide vanes are optimized according to the flow field distribution to ensure smooth flow without dead zones and avoid local overheating of the fairing.

[0069] The present invention carries out temperature field analysis of the internal and external flow fields of the infrared suppressor fairing, realizes the comprehensive optimization of aerodynamic performance and thermal characteristics, and obtains a superior aerodynamic solution for the infrared suppressor fairing.

[0070] Other embodiments of the present disclosure will readily occur to those skilled in the art after considering the specification and practicing the disclosure herein. This application is intended to cover any variations, uses, or adaptations of the present disclosure that follow the general principles of the present disclosure and include common knowledge or customary techniques in the art not disclosed herein. The description and examples are to be considered as exemplary only, with the true scope and spirit of the present disclosure being indicated by the following claims.

[0071] It should be understood that the present disclosure is not limited to the exact structures that have been described above and shown in the drawings, and that various modifications and changes can be made without departing from the scope thereof. The scope of the present disclosure is limited only by the appended claims.

Claims

1. A helicopter infrared suppressor fairing, characterized in that: include: Fairing body, shutter air inlet, square protective net air inlet, secondary exhaust pipe, primary mixing exhaust pipe; The fairing body is viewed from the front as a half ellipse cut along the long axis; the half ellipse is composed of two different upper 1 / 4 ellipse and lower 1 / 4 ellipse; The front of the fairing body is provided with a louver air inlet, the louver plane is crescent-shaped; the angle between the louver plane and the longitudinal symmetry plane is 35°~40°; the louver blades are L-shaped, the angle between the outer edge of the blade and the horizontal plane is 18°~20°, and the angle between the inner edge of the blade and the horizontal plane is 45°~47°; A square protective mesh air inlet is provided at the front of the top of the fairing body; An elliptical secondary exhaust pipe is located at the rear of the top of the fairing body to further cool the high-temperature exhaust gas with the help of the rotor downwash airflow. The primary mixing exhaust pipe is housed within the secondary exhaust pipe, which is connected to the engine outlet. The angle between the secondary exhaust pipe axis and the horizontal plane is 45° to 50°. The gap between the secondary exhaust pipe and the primary mixing exhaust pipe is 40 to 45 mm, and the overlap between the secondary exhaust pipe and the primary mixing exhaust pipe is 50 mm to 150 mm. The rear of the fairing is designed with a streamlined tail cone. The shape of the streamlined tail cone is composed of two parabolas, and the side gradually transitions from an ellipse to a hyperbolic shape. There is a gap between the tail cone and the fuselage skin to facilitate the exhaust of gas inside the infrared suppressor fairing cabin and the intake of external cold air, thereby reducing the cabin temperature. The tail cone aspect ratio γ is 1 to 1.

5. The airflow enters from the louver air inlet and the square protective net air inlet, and is discharged from the gap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe through the high-speed airflow injection effect; The fairing body does not directly contact the first-stage mixing exhaust pipe, and the distance between the fairing body and the first-stage mixing exhaust pipe is maintained at 150mm~200m.

2. A method for aero-thermal optimization design of a helicopter infrared suppressor fairing according to claim 1, characterized in that: include: The infrared suppressor fairing adopts a convex leading edge design to increase the windward area, create a wind-driving effect, and improve the air intake efficiency; The trailing edge of the fairing is tapered to reduce aerodynamic drag; The side of the fairing adopts a multi-section curved surface splicing design, and the curvature of each section of the surface is optimized according to the flow field distribution; An internal airflow channel is designed between the infrared suppressor fairing and the first-stage mixing exhaust pipe to guide the cooling airflow into the first-stage mixing exhaust pipe for mixing; A louver air inlet and a square protective net air inlet are designed on the infrared suppressor fairing, so that air enters from the louver air inlet and the square protective net air inlet, and is discharged from the gap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe through the high-speed airflow injection effect of the first-stage mixing exhaust pipe.

3. The method for aero-thermal optimization design of a helicopter infrared suppressor fairing according to claim 2, characterized in that: The design method is implemented in the following ways: Based on the dimensions of the helicopter engine exhaust port and the internal structure of the infrared suppressor, a parametric model of the infrared suppressor fairing was established. The parameters of the parametric model included the major and minor axes of the elliptical curve of the fairing's side, and the aspect ratio of the tail cone. The purpose of model optimization is to reduce the aerodynamic drag coefficient of the fairing, weaken the aerodynamic interference of the fairing on the horizontal tail, and reduce the surface temperature of the fairing and the average temperature of the tail exhaust. Based on the internal and external flow fields and parameterized models of CFD methods, a simulation analysis model that fully considers the fuselage and engine exhaust system is established; The unstructured tetrahedron grid is used to perform aerodynamic simulation calculations on the simulation analysis model, and the grid is encrypted in the areas where the geometric curvature of the simulation analysis model changes greatly and the flow field changes violently to improve the accuracy of flow field capture; the momentum source method is used to simulate the rotor downwash flow field; the rotor downwash flow field is part of the external flow field; the ROE flux difference splitting format is used for the spatial discretization of the simulation calculation, and the two-equation turbulence model used in the simulation calculation is used. Realizable model, using the DO model for thermal radiation used in simulation calculations; the external flow field uses velocity inlet and pressure outlet boundary conditions, the engine exhaust uses the internal flow field flow inlet boundary, and the object surface uses a no-slip boundary; Based on the aerodynamic simulation results, flow field velocity and temperature field distribution, the aerodynamic and thermal characteristics analysis of the infrared suppressor fairing is carried out.

4. The method for aero-thermal optimization design of a helicopter infrared suppressor fairing according to claim 3 is characterized in that: Conduct aerodynamic and thermal analysis of infrared suppressor fairings, including: By optimizing the three-dimensional shape of the fairing's sides and trailing edge, smoothing the airflow behind the fairing, reducing airflow separation, and minimizing the fairing's outer contour as much as possible, drag is reduced, weakening the adverse interference with the tail aerodynamic surface; To improve the ejection effect, the surface temperature of the infrared suppressor fairing is reduced, and the gap and overlap between the first-stage mixing exhaust pipe and the second-stage exhaust pipe are optimized; The influence of the infrared suppressor fairing shape on the aerodynamic performance and thermal characteristics is comprehensively weighed to determine the aerodynamic scheme of the infrared suppressor fairing.

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