Comprehensive design method for infrared stealth and aerodynamic thermal performance of aero-engine

Abstract

The invention belongs to the field of aero-engines, and particularly relates to a comprehensive design method for infrared stealth and aerodynamic thermal performance of an aero-engine. The method comprises the steps of determining a comprehensive optimization constraint condition and a comprehensive optimization target of infrared stealth and aerodynamic thermal performance of an engine; coupling variables influencing infrared stealth and aerodynamic thermal performance of the engine are determined, aerodynamic thermal cycle parameters of the engine are selected from the coupling variables, and multiple sets of different aerodynamic thermal cycle parameters of the engine are determined according to the change range of the coupling variables; according to the multiple sets of different engine aerodynamic thermodynamic cycle parameters, engine performance parameters and engine forward and backward infrared radiation intensity are calculated; and under the condition that the performance parameters of the engine meet the requirements, the value of the aerodynamic thermodynamic cycle parameters of the engine is determined by taking the minimum forward backward infrared radiation intensity of the engine as the principle, and the infrared stealth and aerodynamic thermodynamic performance comprehensive design of the engine is completed.

Description

Technical Field

[0001] This application belongs to the field of aeroengines, and particularly relates to a comprehensive design method for the infrared stealth and aerodynamic thermal performance of an aeroengine. Background Technique

[0002] There are many coupled design variables for the infrared stealth and aerodynamic thermal performance of an aeroengine. These variables are closely related and there are significant contradictions. The stealth design makes the flow profile of the aeroengine complex and the cooling air volume large, resulting in a large thrust loss.

[0003] The traditional design method first takes performance indicators such as thrust and specific fuel consumption as the goals, determines the aerodynamic thermal cycle parameters and structural parameters of the engine, and then, on this basis, conducts stealth design. By increasing the external and internal compression ratios, increasing the cold air volume of the rear-view visible components, increasing the shielding structure and other measures to improve the stealth ability, but the measures to improve the stealth ability will change the aerodynamic thermal cycle parameters of the engine. For example, the increase in the cooling air volume will increase the mixing loss between the cooling air and the mainstream, and the increase in the shielding structure will lead to an increase in the flow path loss of the mainstream channel, etc., resulting in a significant decrease in performance indicators such as thrust. Therefore, the traditional design method lacks a comprehensive analysis and consideration of the influence of coupled design variables on the infrared stealth and aerodynamic thermal performance of the engine, and cannot achieve the comprehensive optimization design of the infrared stealth and aerodynamic thermal performance of the engine.

[0004] Therefore, there is an urgent need for a technical solution to overcome or mitigate at least one of the above defects of the existing technology. Summary of the Invention

[0005] The purpose of this application is to provide a comprehensive design method for the infrared stealth and aerodynamic thermal performance of an aeroengine to solve at least one problem existing in the prior art.

[0006] The technical solution of this application is as follows: A comprehensive design method for the infrared stealth and aerodynamic thermal performance of an aeroengine, comprising: Step 1, determine the comprehensive optimization constraints and comprehensive optimization objectives for the infrared stealth and aerodynamic thermal performance of the engine. The comprehensive optimization constraints are that the engine performance parameters meet the requirements, and the comprehensive optimization objective is to minimize the infrared radiation intensity in the direct rear direction of the engine; Step 2, determine the coupled variables that affect the infrared stealth and aerodynamic thermal performance of the engine, select the aerodynamic thermal cycle parameters of the engine from the coupled variables, and determine multiple different sets of aerodynamic thermal cycle parameters of the engine according to the change range of the coupled variables; Step 3, calculate the engine performance parameters and the infrared radiation intensity in the direct rear direction of the engine according to multiple different sets of the aerodynamic thermal cycle parameters of the engine; Step 4: On the premise that the engine performance parameters meet the requirements, determine the values of the engine aerodynamic and thermodynamic cycle parameters based on the principle of minimizing the infrared radiation intensity in the direct rear direction of the engine, and complete the comprehensive design of engine infrared stealth and aerodynamic and thermodynamic performance.

[0007] In at least one embodiment of the present application, the engine performance parameters include thrust, specific fuel consumption, fan and compressor pressure ratios, fan and compressor surge margins, and turbine inlet temperature.

[0008] In at least one embodiment of the present application, in Step 2, determine multiple groups of different engine aerodynamic and thermodynamic cycle parameters according to the variation ranges of the coupling variables, including: Determine the variation ranges of the respective coupling variables according to historical data; Determine the preset step sizes of the respective coupling variables; According to the preset step sizes, determine multiple values from the variation ranges of the coupling variables as the values of the corresponding engine aerodynamic and thermodynamic cycle parameters; Combine the values of the respective engine aerodynamic and thermodynamic cycle parameters to obtain multiple groups of different engine aerodynamic and thermodynamic cycle parameters.

[0009] In at least one embodiment of the present application, the engine aerodynamic and thermodynamic cycle parameters include: core outlet gas flow rate, core outlet gas temperature, turbine cooling air flow rate, turbine cooling air temperature, afterburner cooling air flow rate, afterburner cooling air temperature, nozzle cooling air flow rate, nozzle cooling air temperature, nozzle throat area, and nozzle exit area.

[0010] In at least one embodiment of the present application, in Step 3, calculate the engine performance parameters using the overall performance design software Gasturb.

[0011] In at least one embodiment of the present application, in Step 3, the calculation process of the infrared radiation intensity in the direct rear direction of the engine includes: Calculate the average temperature of the visible turbine solid wall surface in the rear view according to the engine aerodynamic and thermodynamic cycle parameters: T 涡轮 =f(W6, T6, W 涡轮冷却气 , T 涡轮冷却气 ); Calculate the average temperature of the visible afterburner solid wall surface in the rear view according to the engine aerodynamic and thermodynamic cycle parameters: T 加力燃烧室 =f(W6, T6, W 加力燃烧室冷却气 , T 加力燃烧室冷却气 ); Calculate the average temperature of the visible nozzle solid wall surface in the rear view according to the engine aerodynamic and thermodynamic cycle parameters: T喷管 = f(W6, T6, W 喷管冷却气 , T 喷管冷却气 ); Among them, T 涡轮 is the average temperature of the solid wall of the visible turbine from the rear view, T 加力燃烧室 is the average temperature of the solid wall of the afterburner visible from the rear view, T 喷管 is the average temperature of the solid wall of the nozzle visible from the rear view, W6 is the gas flow rate at the core engine outlet, T6 is the gas temperature at the core engine outlet, W 涡轮冷却气 is the turbine cooling air flow rate, T 涡轮冷却气 is the turbine cooling air temperature, W 加力燃烧室冷却气 is the afterburner cooling air flow rate, T 加力燃烧室冷却气 is the afterburner cooling air temperature, W 喷管冷却气 is the nozzle cooling air flow rate, T 喷管冷却气 is the nozzle cooling air temperature; Calculate the projected area of the turbine visible area according to the engine aerodynamic and thermodynamic cycle parameters: A 涡轮 = A8 - πR 2 内锥 ; Calculate the projected area of the afterburner visible area according to the engine aerodynamic and thermodynamic cycle parameters: A 加力燃烧室 = πR 2 内锥 ; Calculate the projected area of the nozzle visible area according to the engine aerodynamic and thermodynamic cycle parameters: A 喷管 = A9 - A8; Among them, A 涡轮 is the projected area of the turbine visible area, A 加力燃烧室 is the projected area of the afterburner visible area, A 喷管 is the projected area of the nozzle visible area, A8 is the nozzle throat area, A9 is the nozzle exit area, R 内锥 is the inner cone radius of the afterburner; Calculate the infrared radiation intensity generated by the turbine itself: IR 涡轮 = ε 涡轮 σT 4 涡轮 A 涡轮 ; Calculate the infrared radiation intensity generated by the afterburner itself: IR 加力燃烧室 = ε 加力燃烧室 σT 4 加力燃烧室 A 加力燃烧室 ; Calculate the infrared radiation intensity generated by the nozzle itself: IR 喷管 = ε 喷管 σT 4 喷管 A 喷管 ; Among them, IR 涡轮 is the infrared radiation intensity generated by the turbine itself, IR 加力燃烧室 is the infrared radiation intensity generated by the afterburner itself, IR 喷管 is the infrared radiation intensity generated by the nozzle itself, ε 涡轮 is the infrared emissivity of the solid wall of the turbine, ε 加力燃烧室 is the infrared emissivity of the solid wall of the afterburner, ε 喷管 is the infrared emissivity of the solid wall of the nozzle, and σ is the Stefan-Boltzmann constant; Calculate the infrared radiation intensity directly behind the engine: IR 发动机 = IR 涡轮 + IR 加力燃烧室 + IR 喷管 ; Among them, IR 发动机 is the infrared radiation intensity directly behind the engine.

[0012] In at least one embodiment of the present application, the infrared emissivity ε 涡轮 of the solid wall of the turbine, the infrared emissivity ε 加力燃烧室 of the solid wall of the afterburner, and the infrared emissivity ε 喷管 of the solid wall of the nozzle are determined according to the preliminary design results of component material selection in the engine design stage.

[0013] The invention has at least the following beneficial technical effects: The integrated design method for infrared stealth and aerodynamic thermal performance of the aeroengine in the present application takes the comprehensive optimization of infrared stealth index and aerodynamic thermal performance index as the design goal at the beginning of the engine design stage, analyzes the influence of different coupling design variables on the infrared stealth and aerodynamic thermal performance of the engine, and realizes the integrated optimization design of the infrared stealth and aerodynamic thermal performance of the engine. Description of the Drawings

[0014] Figure 1 is the flowchart of the integrated design method for infrared stealth and aerodynamic thermal performance of the aeroengine in an embodiment of the present application. Detailed Embodiment

[0015] To make the purpose, technical solutions and advantages of the present application more clear, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the accompanying drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals represent the same or similar elements or elements with the same or similar functions from beginning to end. The described embodiments are part of the embodiments of the present application, rather than all of the embodiments. The embodiments described below with reference to the accompanying drawings are exemplary and are intended to explain the present application, and should not be construed as a limitation to the present application. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present application without creative efforts belong to the scope of protection of the present application. The embodiments of the present application will be described in detail below with reference to the accompanying drawings.

[0016] In the description of the present application, it should be understood that the orientation or positional relationships indicated by the terms "center", "longitudinal", "lateral", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc. are based on the orientation or positional relationships shown in the drawings, and are only for the convenience of describing the present application 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 should not be construed as a limitation to the scope of protection of the present application.

[0017] The following will further describe the present application in detail with reference to the attached Figure 1 drawings.

[0018] The present application provides a comprehensive design method for the infrared stealth and aerodynamic thermal performance of an aeroengine, including the following steps: Step 1: Determine the comprehensive optimization constraints and comprehensive optimization objectives for the infrared stealth and aerodynamic thermal performance of the engine. The comprehensive optimization constraints are that the engine performance parameters meet the requirements, and the comprehensive optimization objective is to minimize the infrared radiation intensity in the directly rearward direction of the engine; Step 2: Determine the coupling variables that affect the infrared stealth and aerodynamic thermal performance of the engine, select the engine aerodynamic thermal cycle parameters from the coupling variables, and determine multiple groups of different engine aerodynamic thermal cycle parameters according to the variation ranges of the coupling variables; Step 3: Calculate the engine performance parameters and the infrared radiation intensity in the directly rearward direction of the engine according to multiple groups of different engine aerodynamic thermal cycle parameters; Step 4: When the engine performance parameters meet the requirements, determine the values of the engine aerodynamic thermal cycle parameters based on the principle of minimizing the infrared radiation intensity in the directly rearward direction of the engine, and complete the comprehensive design of the infrared stealth and aerodynamic thermal performance of the engine.

[0019] The comprehensive design method for infrared stealth and aerodynamic thermal performance of an aeroengine in this application. First, in step one, based on the comprehensive optimization constraint conditions where the engine performance parameters meet the requirements, the minimum infrared radiation intensity in the direct rear direction of the engine is added as the comprehensive optimization objective. In the preferred embodiment of this application, the engine performance parameters mainly include thrust, specific fuel consumption, fan and compressor pressure ratios, fan and compressor surge margins, turbine inlet temperature, etc. The infrared radiation in the direct rear direction of the engine mainly consists of the thermal wall surface radiation of the rear-view visible components, jet radiation, and reflected radiation of the visible area projection area. Since the jet radiation and the reflected radiation of the visible area projection area account for a relatively small proportion and can be ignored in the calculation, only the thermal wall surface radiation of the rear-view visible components is considered in this application for the one-dimensional rapid calculation of the infrared radiation in the direct rear direction of the engine.

[0020] The comprehensive design method for infrared stealth and aerodynamic thermal performance of an aeroengine in this application. Second, in step two, determine the coupling variables that affect the infrared stealth and aerodynamic thermal performance of the engine, and select the engine aerodynamic thermal cycle parameters from the coupling variables. The engine aerodynamic thermal cycle parameters mainly include: core outlet gas flow rate, core outlet gas temperature, turbine cooling air flow rate, turbine cooling air temperature, afterburner cooling air flow rate, afterburner cooling air temperature, nozzle cooling air flow rate, nozzle cooling air temperature, nozzle throat area, nozzle exit area, etc.

[0021] In the preferred embodiment of this application, the process of determining multiple sets of different engine aerodynamic thermal cycle parameters according to the change range of the coupling variables is as follows: Determine the change range of each coupling variable according to historical data; Determine the preset step size of each coupling variable; According to the preset step size, determine multiple values from the change range of the coupling variables as the values of the corresponding engine aerodynamic thermal cycle parameters; Combine the values of each engine aerodynamic thermal cycle parameter to obtain multiple sets of different engine aerodynamic thermal cycle parameters.

[0022] In the historical data that meet the optimization constraint conditions of the conventional aeroengine aerodynamic thermal cycle parameters, determine the change range of the coupling variables that affect the infrared stealth and aerodynamic thermal performance of the engine. Adjust the values of each engine aerodynamic thermal cycle parameter one by one at a certain step size within the change range to form multiple sets of different aeroengine thermal cycle parameters for calculating the engine performance parameters and the infrared radiation intensity in the direct rear direction of the engine.

[0023] In the comprehensive design method for infrared stealth and aerodynamic thermal performance of an aeroengine according to the present application, in step 3, after obtaining the values of various aeroengine aerodynamic thermal cycle parameters, the aeroengine performance parameters and the infrared radiation intensity in the direct rearward direction of the aeroengine are calculated. In a preferred embodiment of the present application, the values of various aeroengine aerodynamic thermal cycle parameters are input into the overall performance design software Gasturb, and the aeroengine performance parameters are calculated through the overall performance design software Gasturb.

[0024] In this embodiment, the calculation process of the infrared radiation intensity in the direct rearward direction of the aeroengine includes: Calculating the average temperature of the visible solid wall surface of the rear-view turbine according to the aeroengine aerodynamic thermal cycle parameters: T 涡轮 =f(W6, T6, W 涡轮冷却气 , T 涡轮冷却气 ); Calculating the average temperature of the visible solid wall surface of the afterburner in the rear view according to the aeroengine aerodynamic thermal cycle parameters: T 加力燃烧室 =f(W6, T6, W 加力燃烧室冷却气 , T 加力燃烧室冷却气 ); Calculating the average temperature of the visible solid wall surface of the nozzle in the rear view according to the aeroengine aerodynamic thermal cycle parameters: T 喷管 =f(W6, T6, W 喷管冷却气 , T 喷管冷却气 ); Wherein, T 涡轮 is the average temperature of the visible solid wall surface of the rear-view turbine, T 加力燃烧室 is the average temperature of the visible solid wall surface of the afterburner in the rear view, T 喷管 is the average temperature of the visible solid wall surface of the nozzle in the rear view, W6 is the gas flow rate at the outlet of the core flow, T6 is the gas temperature at the outlet of the core flow, W 涡轮冷却气 is the turbine cooling air flow rate, T 涡轮冷却气 is the turbine cooling air temperature, W 加力燃烧室冷却气 is the afterburner cooling air flow rate, T 加力燃烧室冷却气 is the afterburner cooling air temperature, W 喷管冷却气 is the nozzle cooling air flow rate, T 喷管冷却气 is the nozzle cooling air temperature; Calculating the projected area of the visible area of the turbine according to the aeroengine aerodynamic thermal cycle parameters: A 涡轮 =A8 - πR 2 内锥 ; Calculating the projected area of the visible area of the afterburner according to the aeroengine aerodynamic thermal cycle parameters: A 加力燃烧室 =πR 2内锥 ; Calculate the projected area of the nozzle visible region based on the engine aerodynamic and thermodynamic cycle parameters: A 喷管 = A9 - A8; where, A 涡轮 is the projected area of the turbine visible region, A 加力燃烧室 is the projected area of the afterburner visible region, A 喷管 is the projected area of the nozzle visible region, A8 is the nozzle throat area, A9 is the nozzle exit area, R 内锥 is the inner cone radius of the afterburner; Calculate the infrared radiation intensity generated by the turbine itself: IR 涡轮 = ε 涡轮 σT 4 涡轮 A 涡轮 ; Calculate the infrared radiation intensity generated by the afterburner itself: IR 加力燃烧室 = ε 加力燃烧室 σT 4 加力燃烧室 A 加力燃烧室 ; Calculate the infrared radiation intensity generated by the nozzle itself: IR 喷管 = ε 喷管 σT 4 喷管 A 喷管 ; where, IR 涡轮 is the infrared radiation intensity generated by the turbine itself, IR 加力燃烧室 is the infrared radiation intensity generated by the afterburner itself, IR 喷管 is the infrared radiation intensity generated by the nozzle itself, ε 涡轮 is the infrared emissivity of the turbine solid wall, ε 加力燃烧室 is the infrared emissivity of the afterburner solid wall, ε 喷管 is the infrared emissivity of the nozzle solid wall, σ is the Stefan - Boltzmann constant; Calculate the forward - facing infrared radiation intensity of the engine: IR 发动机 = IR 涡轮 + IR 加力燃烧室 + IR 喷管 ; where, IR 发动机 is the forward - facing infrared radiation intensity of the engine.

[0025] Determine the infrared emissivity ε of the turbine solid wall according to the preliminary design results of component material selection in the engine scheme design stage涡轮 , the infrared emissivity ε of the solid wall of the afterburner 加力燃烧室 and the infrared emissivity ε of the solid wall of the nozzle 喷管 .

[0026] For the comprehensive design method of infrared stealth and aerodynamic thermal performance of the aeroengine of the present application, finally, in step four, according to the engine performance parameters and the calculation results of the infrared radiation intensity in the directly backward direction of the engine, when the engine performance parameters meet the requirements, the values of the various engine aerodynamic thermal cycle parameters corresponding to the minimum infrared radiation intensity in the directly backward direction of the engine are obtained, and the comprehensive design of the infrared stealth and aerodynamic thermal performance of the engine is completed according to these engine aerodynamic thermal cycle parameters that affect the infrared stealth and aerodynamic thermal performance of the engine.

[0027] The comprehensive design method of infrared stealth and aerodynamic thermal performance of the aeroengine of the present application is applicable to the comprehensive design of infrared stealth and aerodynamic thermal performance of small bypass ratio afterburning turbofan engines, can realize the comprehensive optimization design of infrared stealth and aerodynamic thermal performance of the engine at the engine scheme design stage, while greatly improving the infrared stealth performance of the engine, minimizes the influence on engine performance indicators such as thrust and specific fuel consumption, and provides strong support for the comprehensive design of the engine.

[0028] The above is only the specific implementation manner of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed in the present application should be covered by the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claimed rights.

Claims

1. An integrated design method for infrared stealth and aerodynamic thermal performance of an aeroengine, characterized in that Including: Step 1: Determine the comprehensive optimization constraints and comprehensive optimization objectives for the infrared stealth and aerodynamic thermal performance of the engine. The comprehensive optimization constraints are that the engine performance parameters meet the requirements, and the comprehensive optimization objective is to minimize the infrared radiation intensity in the directly backward direction of the engine; Step 2: Determine the coupling variables that affect the infrared stealth and aerodynamic thermal performance of the engine, select the engine aerodynamic thermal cycle parameters from the coupling variables, and determine multiple groups of different engine aerodynamic thermal cycle parameters according to the variation ranges of the coupling variables; Step 3: Calculate the engine performance parameters and the infrared radiation intensity in the directly backward direction of the engine according to multiple groups of different engine aerodynamic thermal cycle parameters; Step 4: When the engine performance parameters meet the requirements, determine the values of the engine aerodynamic thermal cycle parameters based on the principle of minimizing the infrared radiation intensity in the directly backward direction of the engine, and complete the comprehensive design of the infrared stealth and aerodynamic thermal performance of the engine.

2. The integrated design method for infrared stealth and aerodynamic thermal performance of an aeroengine according to claim 1, wherein The engine performance parameters include thrust, specific fuel consumption, fan and compressor pressure ratios, fan and compressor surge margins, and turbine inlet temperature.

3. The integrated design method for infrared stealth and aerodynamic thermal performance of an aeroengine according to claim 2, characterized in that, In Step 2, determining multiple groups of different engine aerodynamic thermal cycle parameters according to the variation ranges of the coupling variables includes: Determine the variation ranges of each of the coupling variables according to historical data; Determine the preset step sizes of each of the coupling variables; According to the preset step sizes, determine multiple values from the variation ranges of the coupling variables as the values of the corresponding engine aerodynamic thermal cycle parameters; Combine the values of each of the engine aerodynamic thermal cycle parameters to obtain multiple groups of different engine aerodynamic thermal cycle parameters.

4. The integrated design method for infrared stealth and aerodynamic thermal performance of an aeroengine according to claim 3, wherein The engine aerodynamic thermal cycle parameters include: core outlet gas flow rate, core outlet gas temperature, turbine cooling air flow rate, turbine cooling air temperature, afterburner cooling air flow rate, afterburner cooling air temperature, nozzle cooling air flow rate, nozzle cooling air temperature, nozzle throat area, and nozzle exit area.

5. The integrated design method for infrared stealth and aerodynamic thermal performance of an aeroengine according to claim 4, characterized in that In Step 3, calculate the engine performance parameters using the overall performance design software Gasturb.

6. The integrated design method for infrared stealth and aerodynamic thermal performance of an aeroengine according to claim 5, characterized in that, In Step 3, the calculation process of the infrared radiation intensity in the directly backward direction of the engine includes: Calculate the average temperature of the visible solid wall surface of the turbine in the rear view according to the engine aerodynamic thermal cycle parameters; T 涡轮 = f(W6, T6, W 涡轮冷却气 , T 涡轮冷却气 ); Calculate the average temperature of the visible solid wall surface of the afterburner in the rear view according to the engine aerodynamic thermal cycle parameters; T 加力燃烧室 = f(W6, T6, W 加力燃烧室冷却气 , T 加力燃烧室冷却气 ); Calculate the average temperature of the visible solid wall surface of the nozzle in the rear view according to the engine aerodynamic thermal cycle parameters; T 喷管 = f(W6, T6, W 喷管冷却气 , T 喷管冷却气 ); Among them, T 涡轮 is the average temperature of the solid wall of the turbine visible from the rear, T 加力燃烧室 is the average temperature of the solid wall of the afterburner visible from the rear, T 喷管 is the average temperature of the solid wall of the nozzle visible from the rear, W6 is the gas flow rate at the outlet of the core flow path, T6 is the gas temperature at the outlet of the core flow path, W 涡轮冷却气 is the cooling air flow rate of the turbine, T 涡轮冷却气 is the cooling air temperature of the turbine, W 加力燃烧室冷却气 is the cooling air flow rate of the afterburner, T 加力燃烧室冷却气 is the cooling air temperature of the afterburner, W 喷管冷却气 is the cooling air flow rate of the nozzle, T 喷管冷却气 is the cooling air temperature of the nozzle; Calculate the projected area of the visible region of the turbine according to the engine aerodynamic thermal cycle parameters; A 涡轮 = A8 - πR 2 内锥 ; Calculate the projected area of the visible region of the afterburner according to the engine aerodynamic thermal cycle parameters; A 加力燃烧室 =πR 2 内锥 ; Calculate the projected area of the visible region of the nozzle according to the engine aerodynamic thermal cycle parameters; A 喷管 = A9 - A8; Among them, A 涡轮 is the projected area of the turbine visible area, A 加力燃烧室 is the projected area of the afterburner visible area, A 喷管 is the projected area of the nozzle visible area, A8 is the nozzle throat area, A9 is the nozzle exit area, R 内锥 is the inner cone radius of the afterburner; Calculate the infrared radiation intensity generated by the turbine itself; IR 涡轮 = ε 涡轮 σT 4 涡轮 A 涡轮 ; Calculate the infrared radiation intensity generated by the afterburner itself; IR 加力燃烧室 = ε 加力燃烧室 σT 4 加力燃烧室 A 加力燃烧室 ; Calculate the infrared radiation intensity generated by the nozzle itself; IR 喷管 = ε 喷管 σT 4 喷管 A 喷管 ; Among them, IR 涡轮 is the infrared radiation intensity generated by the turbine itself, IR 加力燃烧室 is the infrared radiation intensity generated by the afterburner itself, IR 喷管 is the infrared radiation intensity generated by the nozzle itself, ε 涡轮 is the infrared emissivity of the solid wall of the turbine, ε 加力燃烧室 is the infrared emissivity of the solid wall of the afterburner, ε 喷管 is the infrared emissivity of the solid wall of the nozzle, and σ is the Stefan-Boltzmann constant; Calculate the infrared radiation intensity in the directly backward direction of the engine. IR 发动机 = IR 涡轮 + IR 加力燃烧室 + IR 喷管 ; Among them, IR 发动机 is the infrared radiation intensity directly behind the engine.

7. The integrated design method for infrared stealth and aerodynamic thermal performance of an aeroengine according to claim 6, characterized in that, Determine the infrared emissivity ε of the turbine solid wall surface according to the preliminary design results of component material selection in the engine scheme design stage 涡轮 , the infrared emissivity ε of the solid wall surface of the afterburner 加力燃烧室 and the infrared emissivity ε of the solid wall surface of the nozzle 喷管 .

Images