Forward design method for thermal protection structure of engine housing bullet wall cable

Abstract

The invention discloses a forward design method for the thermal protection structure of the bullet wall cable of the engine housing, including: taking the distribution of the overheat ablation area on the inner wall of the engine during the active flight section of the aircraft and the distribution of the aerodynamic heating on the outer wall of the engine during the whole flight process of the aircraft as the design constraints of the bullet wall cable, and determining the circumferential layout position of the bullet wall cable on the engine housing; adopting the method of arranging the bullet wall cable on the inner wall of the engine housing, and wrapping the bullet wall cable in the thermal protection material insulation layer to form a thermal protection structure; wherein, the outer wall of the insulation layer is attached to the inner wall of the engine housing, and the inner wall of the insulation layer is attached to the outer wall of the engine charge. The method of the invention effectively solves the problems of more scheme iteration times, longer design cycle and increased design cost caused by the traditional engineering experience-based repeated iterative design.

Description

Technical Field

[0001] The present invention relates to the field of thermal protection for aircraft engine, and particularly to a forward design method for the thermal protection structure of the cable on the bullet wall of the engine housing. Background Art

[0002] For aircraft, there are mainly two types of commonly used cables on the bullet wall of the engine housing at present. One type is to arrange the bullet wall cable outside the engine housing, and use a cable heat shield on the outer wall of the engine housing to protect the cable on the outer wall of the engine housing; the other type is to arrange the bullet wall cable inside the engine housing and use an insulating layer to protect the cable on the inner wall.

[0003] For the solution of arranging the bullet wall cable outside the engine housing and using a cable heat shield to protect the cable, since the heat shield forms a local protrusion in the circumferential direction of the engine housing, during the flight of the aircraft, the aerodynamic heating at the protruding position of the heat shield is enhanced, the local electromagnetic environment deteriorates, and the aerodynamic shear force at the connection between the heat shield and the outer wall of the engine housing is also relatively strong. The existing problems are as follows: In terms of the aerodynamic shape of the aircraft, the structure of the cable heat shield on the outer wall will cause a protrusion in the circumferential direction of the outer wall of the engine housing, which is not conducive to maintaining the overall aerodynamic shape of the engine compartment of the aircraft, and may have an adverse impact on the aerodynamic performance of the aircraft; in terms of structural design and manufacturing process, for the cable heat shield, it is necessary to have strong thermal protection ability and electromagnetic interference prevention ability, and the connection between the engine and the engine housing needs to be strengthened; in terms of the adaptability to the launch platform, for the structural solution of the cable heat shield on the outer wall, the local protrusion caused by it may pose a risk that it is difficult to match the interface size of the launch platform; in terms of storage and transportation, the protruding cable heat shield on the outer wall has a risk of being knocked during storage and transportation, and additional buffer protection treatment is required, which has higher requirements for the impact resistance of the storage box.

[0004] For the solution of arranging the bullet wall cable inside the engine housing and using an insulating layer to protect the cable on the inner wall, this type of solution has high requirements for the heat insulation effect of the insulating layer and the specific arrangement position of the cable in the circumferential direction of the inner wall of the engine housing. Especially in terms of the selection of the arrangement direction of the cable on the inner wall of the engine housing, there is no clear design specification, specific requirement or guidance on how to select a suitable cable arrangement position according to the thermal environment distribution conditions of the inner and outer walls of the engine during the flight of the aircraft. Therefore, in engineering practical applications, it is often necessary to design based on engineering experience, which may lead to problems such as more iterations of the solution, repeated design, resulting in a longer design cycle and an increase in design cost, and is not conducive to the establishment of a forward design method. Summary of the Invention

[0005] The purpose of the present invention is to provide a forward design method for the thermal protection structure of the cable on the bullet wall of the engine housing to overcome the problems existing in the wiring process of the existing cable on the bullet wall of the engine housing.

[0006] To achieve the above tasks, the present invention adopts the following technical solutions:

[0007] A forward design method for the thermal protection structure of the bullet wall cable of an engine housing, comprising:

[0008] Taking the distribution of the overheat ablation area on the inner wall of the engine during the active flight section of the aircraft and the distribution of the aerodynamic heating on the outer wall of the engine throughout the flight process of the aircraft as the design constraints of the bullet wall cable, and determining the circumferential arrangement position of the bullet wall cable on the engine housing;

[0009] Adopting the method of arranging the bullet wall cable on the inner wall of the engine housing, and wrapping the bullet wall cable in the thermal insulation layer of the thermal protection material to form a thermal protection structure; wherein, the outer wall of the thermal insulation layer is attached to the inner wall of the engine housing, and the inner wall of the thermal insulation layer is attached to the outer wall of the engine charge.

[0010] Further, the specific principle for determining the circumferential arrangement position of the bullet wall cable on the engine housing is:

[0011] Avoiding the overheat ablation area on the inner wall of the engine and the affected adjacent areas during the active flight section of the aircraft:

[0012] ;

[0013] Wherein, represents existence, represents the circumferential arrangement angle of the bullet wall cable on the engine housing, k represents the boundary margin of the overheat ablation area on the inner wall of the engine, α represents the circumferential distribution angle of the overheat ablation area on the inner wall of the engine on the engine housing, , represents the minimum and maximum azimuth angles of the active section of the aircraft;

[0014] Among the four regions of the outer wall of the engine housing divided by counterclockwise rotation according to the Y axis and the Z axis in the body coordinate system of the bullet: the windward side, the right side, the leeward side, and the left side, selecting the region with the minimum average heat flux density for arrangement;

[0015] The circumferential position distribution of the bullet wall cable on the inner wall of the engine housing should not interfere with other components inside the engine.

[0016] Further, the specific determination method for the overheat ablation area on the inner wall of the engine and the affected adjacent areas during the active flight section of the aircraft is:

[0017] According to the magnitude and direction of the flight lateral overload in the active section of the aircraft flight, obtain the circumferential position distribution of the overload ablation area on the inner wall of the engine within the inner wall of the engine casing;

[0018] Combined with the flight ballistic data of the aircraft, perform ballistic calculation according to the principles of aircraft flight mechanics. Through the Y lateral overload and Z longitudinal overload synthesis in the body coordinate system of the projectile, the flight lateral overload can be obtained;

[0019] Take the position on the inner wall of the engine casing corresponding to the reverse side of the azimuth angle of the flight lateral overload as the overload ablation area on the inner wall of the engine;

[0020] Record the maximum flight time of the active section of the aircraft as , when the flight lateral overload varies with time t , the azimuth angle also varies with time t , and the variation range is , where ; represents the minimum and maximum azimuth angles of the active section of the aircraft;

[0021] Define the angle of the circumferential distribution of the overload ablation area on the inner wall of the engine as α , then there is:

[0022] ;

[0023] Define the boundary margin of the overload ablation area on the inner wall of the engine as k , where ; the overload ablation area on the inner wall of the engine and the affected adjacent areas S GZ are defined as:

[0024] .

[0025] Furthermore, the determination method of the four regions on the outer wall of the engine casing is as follows:

[0026] Based on the body coordinate system of the aircraft projectile, in the YOZ plane, rotate the Y axis and Z axis of the body coordinate system counterclockwise by 45°. Then, on the outer wall of the engine casing in the YOZ plane, it will be divided into four regions with a central angle of 90° by the rotated Y axis and Z axis. The area directly below is the windward side, the right side area is the right side, the area directly above is the leeward side, and the left side area is the left side.

[0027] Further, the method for determining the average heat flux density is as follows:

[0028] For the aerodynamic heating on the outer wall of the engine during the entire flight process of the aircraft, it is also combined with the flight trajectory data, and through computational fluid dynamics methods, the variation law of the surface heat flux density on the outer wall of the engine housing with time is obtained;

[0029] The variation law is partitioned according to the four divided regions to obtain the heat flux density of each region; then the heat flux density of each region is spatially averaged respectively to obtain the average heat flux density varying with time for each region.

[0030] Further, in the protection structure, the thickness from the inner wall of the heat insulation layer to the cable on the missile wall d in is determined as follows:

[0031] Select the temperature in the region close to the inner wall of the heat insulation layer inside the engine housing T in ( t ) as the temperature boundary condition for the inner wall region of the heat insulation layer, where the time , is the maximum flight time of the active section of the aircraft, define the allowable working temperature of the cable on the missile wall as T a , and the initial temperature is selected as the ambient temperature T 0;

[0032] Establish a first heat transfer finite element model to obtain the density, specific heat capacity, and thermal conductivity of the heat insulation layer material; only consider heat conduction for the heat transfer process from the inner wall of the heat insulation layer to the cable on the missile wall; through finite element calculation, for d in , within the time range of 0 to t act , when the highest temperature T inXL_Max on the side of the cable on the missile wall close to the inner wall of the heat insulation layer does not exceed the allowable working temperature T a , d in the design of

[0033] Further, in the protection structure, the thickness from the outer wall of the heat insulation layer to the cable on the missile wall d out is determined as follows:

[0034] Obtain the average heat flux density of four regions on the outer wall of the engine casing, the material parameters of the insulation layer, the material density, specific heat capacity, and thermal conductivity of the engine casing; establish a second heat transfer finite element model. The heat transfer process from the outer wall of the engine casing to the outer wall of the insulation layer and then from the outer wall of the insulation layer to the cable on the casing wall only considers heat conduction; through finite element calculation, for d out , within the time range of 0 to t all , when the maximum temperature on the side of the cable on the casing wall close to the outer wall of the insulation layer T outXL_Max does not exceed the allowable working temperature T a , d out the design meets the thermal protection requirements; t all represents the maximum time during the entire flight process of the aircraft.

[0035] Furthermore, the method further includes: locally arranging a thermal protection coating on the outer wall of the engine casing; the circumferential distribution range of the thermal protection coating on the outer wall of the engine casing not only needs to completely cover the circumferential distribution range of the cable on the casing wall inside the engine casing, but also extends circumferentially at the circumferential boundary position of the thermal protection coating.

[0036] Furthermore, define the circumferential distribution angle of the thermal protection coating on the outer wall of the engine casing as φ tc , and define the boundary margin of the cable wiring position on the casing wall of the engine casing as l , where ; then there is:

[0037] ;

[0038] wherein, β represents the circumferential arrangement angle of the cable on the casing wall of the engine casing ; , respectively represent the maximum and minimum values of the circumferential arrangement angle of the cable on the casing wall of the engine casing.

[0039] Furthermore, for the thickness d tc of the thermal protection coating, it is obtained by establishing a third heat transfer finite element model;

[0040] During modeling, in addition to the material parameters of the insulation layer and the engine housing, it is also necessary to clarify the density, specific heat capacity, and thermal conductivity of the thermal protection coating; the average heat flux density of the four regions on the outer wall of the engine housing is applied to the outer surface of the thermal protection coating; the heat transfer process from the thermal protection coating to the outer wall of the engine housing, from the outer wall of the engine housing to the outer wall of the insulation layer, and then from the outer wall of the insulation layer to the missile wall cable only considers heat conduction; through finite element calculation, for the thickness of the thermal protection coating d tc , within the range of 0 to t all time, when the maximum temperature on the side of the missile wall cable close to the thermal protection coating after setting the thermal protection coating T tcXL_Max does not exceed the allowable working temperature T a , d tc the design of t all meets the thermal protection requirements;

[0041] Compared with the prior art, the present invention has the following technical features:

[0042] 1. Compared with the structural solution of arranging the missile wall cable outside the engine housing and using a heat shield to protect the outer wall cable on the outer wall of the engine housing, the design of the present invention has the characteristics of maintaining the aerodynamic shape of the engine housing of the aircraft, light structural weight, good processability, low cost, easy installation, strong adaptability to the launch platform, and good storage and transportation performance.

[0043] 2. Compared with the solution of arranging the missile wall cable inside the engine housing and using an insulation layer to protect the inner wall cable, in the present invention, a thermal protection coating material with a relatively thin thickness is also used on the local outer wall of the engine housing corresponding to the position of the missile wall cable, further enhancing the thermal protection ability of the missile wall cable during the entire flight process of the aircraft.

[0044] 3. The forward design method of structural thermal protection of the present invention, by establishing a finite element simulation model, taking the internal and external wall thermal environments of the engine housing as design inputs, clarifying the boundary conditions of the cable thermal protection structure, calculating and analyzing the relevant parameters of the structure, and optimizing the design of the thermal protection structure, and the relevant structural parameters are controllable and can be quickly iterated in the established finite element model. This method can solve the problems of a large number of scheme iterations, long design cycles, and increased design costs caused by traditional engineering experience-based repeated iterative designs. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] Figure 1 is a schematic diagram of the division of four regions in the YOZ plane of the aircraft body coordinate system.

[0046] Figure 2 Schematic diagram of the thermal protection structure in an embodiment of the present invention.

[0047] Explanation of reference numerals: 1 engine housing, 2 thermal insulation layer, 3 engine charge, 4 cable on the missile wall, 5 thermal protection coating. Detailed implementation manners

[0048] Based on the scheme of arranging the cable on the missile wall inside the engine housing and using the thermal insulation layer to protect the inner wall cable, on the basis of its advantages of being able to maintain the aerodynamic shape of the aircraft engine housing, having a light structural weight, good processability, low cost, and easy installation, the present invention provides a forward design method for the cable on the missile wall of the engine housing. Based on the idea of forward design, the thermal environment on the inner and outer walls of the engine housing during the flight of the aircraft is used as the design input. Through the establishment of a finite element model for simulation analysis, the relevant parameters of the cable thermal protection structure are obtained, and the thermal protection structure is optimized and designed, so as to solve the problems of more iteration times of the scheme, longer design cycle, and increased design cost caused by the traditional engineering experience-based repeated iterative design.

[0049] A forward design method for the thermal protection structure of the cable on the missile wall of the engine housing provided by the present invention includes the following steps:

[0050] Step 1, taking the distribution of the overloading ablation area on the inner wall of the engine during the active flight section of the aircraft and the distribution of the aerodynamic heating on the outer wall of the engine during the whole flight process of the aircraft as the design constraints for the cable on the missile wall, and determining the circumferential arrangement position of the cable on the missile wall in the engine housing.

[0051] Generally speaking, the circumferential positions of the most severely overloaded ablation area on the inner wall of the engine during the active flight section of the aircraft and the aerodynamic heating area on the outer wall of the engine during the whole flight process of the aircraft are different. Therefore, the circumferential arrangement position of the cable on the missile wall should be avoided as much as possible from the above two high-temperature areas.

[0052] (1.1) Overloading ablation area on the inner wall of the engine during the active flight section of the aircraft.

[0053] According to the magnitude and direction of the lateral overload during the active flight section of the aircraft, the circumferential position distribution of the overloading ablation area on the inner wall of the engine in the engine housing can be obtained; the position of the overloading ablation area on the inner wall of the engine is generally on the side opposite to the direction of the lateral overload. The lateral overload can be combined with the flight ballistic data of the aircraft and the ballistic solution can be carried out according to the principles of aircraft flight mechanics. Through the Y lateral overload n y and Z lateral overload n z synthesis, the lateral overload during flight can be obtainedn yz Size:

[0054] ;

[0055] Among them, the projectile coordinate system takes the axial direction of the aircraft as the X axis, and on the cross-section of the aircraft, takes the axis position of the aircraft as the O point to establish the Y axis and the Z axis, as shown in Figure 1 .

[0056] Take the flight lateral overload n yz azimuth angle at the position on the inner wall of the engine housing corresponding to the reverse side as the overheat ablation area of the inner wall of the engine; among them, the azimuth angle is expressed as follows:

[0057] ;

[0058] Record the maximum flight time of the active section of the aircraft as t act , and the maximum time of the entire flight process is t all . When the flight lateral overload varies with time t , the azimuth angle also varies with time t , and its variation range is , which is the circumferential distribution of the overheat ablation area of the inner wall of the engine, where ; represents the minimum and maximum azimuth angles of the active section of the aircraft.

[0059] Define the angular size of the circumferential distribution of the overheat ablation area of the inner wall of the engine as α , then there is:

[0060] ;

[0061] Define the boundary margin of the overheat ablation area of the inner wall of the engine as k , where ; Considering the influence of the boundary of the overheat ablation area of the inner wall of the engine on the adjacent area, in order to strengthen the safety margin of the design, the layout area of the cable on the projectile wall should also avoid the adjacent area; the overheat ablation area of the inner wall of the engine and the affected adjacent area S GZ is defined as:

[0062] ;

[0063] For example, when k takes 0.2, it means that the influence range of the overloaded ablation area on the inner wall of the engine on the adjacent area of the unilateral boundary accounts for 20% of the circumferential distribution angle range of the ablation overloaded area, and the interval that the circumferential arrangement position of the cable on the missile wall needs to avoid is S GZ ; k The larger the value, the less likely the circumferential arrangement position of the cable on the missile wall is affected by the overloaded ablation area of the engine, and the greater the designed safety margin.

[0064] (1.2) Aerodynamic heating of the outer wall of the engine during the whole flight process of the aircraft.

[0065] The aerodynamic heating of the outer wall of the engine during the whole flight process of the aircraft is also combined with the flight trajectory data, and the surface heat flux density of the outer wall of the engine housing is obtained through the computational fluid dynamics (CFD) method q w with time t variation law q w ( t ).

[0066] For the convenience of understanding, based on the missile body coordinate system of the aircraft, in the YOZ plane, the Y axis and Z axis of the missile body coordinate system are rotated counterclockwise by 45°, then in the YOZ plane, the outer wall of the engine housing will be divided into four regions with a central angle of 90° by the rotated Y axis and Z axis. Among them, the area directly below is the windward side S wind , the right area is the right side S Rside , the area directly above is the leeward side S lee , and the left area is the left side S Lside , as Figure 1 shown.

[0067] For the convenience of calculation, the surface heat flux density q w of the outer wall of the engine housing with time t variation law q w ( t ) is partitioned according to the above four regions to obtain the heat flux density of each region; then the heat flux density of each region is spatially averaged respectively to obtain the average heat flux density of each region varying with time t , that is, the average heat flux density of the windward side q w_wind (t ) Average heat flux density of the right side q w_Rside ( t ) Average heat flux density of the leeward side q w_lee ( t ) Average heat flux density of the left side q w_ Lside ( t )

[0068] During the flight of the aircraft, for most cases, the law of the average heat flux density is generally: q w_wind ( t ) > q w_Rside ( t ) ≈ q w_ Lside ( t ) > q w_lee ( t ) However, in actual applications, it should be comprehensively analyzed in combination with the ballistic flight data. For example, when the projectile undergoes axial roll, pitch motion, and yaw motion during the flight, the above law may no longer apply; the circumferential arrangement position of the cable on the projectile wall should be selected from the four regions with the smallest surface heat flux density.

[0069] When designing the circumferential arrangement position of the cable on the projectile wall, in order to improve the overall thermal protection ability of the cable in the radial direction of the engine housing, the cable on the projectile wall is arranged flatly in the circumferential direction; define the maximum and minimum values of the circumferential arrangement angle of the cable on the projectile wall in the engine housing as 、 , then the angle of the circumferential center line position of the cable on the projectile wall is , where = 1 / 2( + ) Define the interval range of the circumferential distribution angle of the cable on the projectile wall as S XL = , , for the circumferential arrangement angle of the cable on the projectile wall in the engine housing, there is always .

[0070] Define the circumferential arrangement angle of the cable on the projectile wall in the engine housing as β , then there is:

[0071] .

[0072] In summary, the design principle for the circumferential arrangement position of the cable on the projectile wall in the engine housing is:

[0073] (1) Avoid the engine inner wall overheat ablation area and the affected adjacent areas during the active flight section of the aircraft, i.e.:

[0074] .

[0075] (2) Among the four areas of the outer wall of the engine housing divided counterclockwise according to the Y axis and the Z axis of the body coordinate system: the windward side, the right side, the leeward side, and the left side, select the area with the minimum average heat flux density for layout.

[0076] (3) Considering the installation convenience of the cable on the missile wall, its circumferential position distribution on the inner wall of the engine housing should not interfere with other components inside the engine.

[0077] Step 2, adopt the method of arranging the missile wall cable on the inner wall of the engine housing, and wrap the missile wall cable in the thermal protection material insulation layer to form a thermal protection structure to maintain the circumferential consistency of the aerodynamic shape of the engine housing; among them, the outer wall of the insulation layer fits with the inner wall of the engine housing, and the inner wall of the insulation layer fits with the outer wall of the engine charge area.

[0078] The thickness from the inner wall of the insulation layer to the missile wall cable d in is obtained by taking the temperature at the engine charge area during the active flight section of the aircraft as the calculation input and the temperature of the missile wall cable not exceeding the allowable working temperature as the constraint design; the thickness from the outer wall of the insulation layer to the missile wall cable d out is obtained by taking the heat flux density on the outer wall of the engine housing during the whole flight process of the aircraft as the calculation input and the temperature of the missile wall cable not exceeding the allowable working temperature as the constraint design, and the specific description is as follows:

[0079] For the thickness from the inner wall of the insulation layer to the missile wall cable d in , during the continuous combustion of the engine charge in the active flight section of the aircraft, to simplify the calculation model, select the temperature T in ( t ) of the area close to the inner wall of the insulation layer inside the engine housing as the temperature boundary condition of the inner wall area of the insulation layer, where ; this temperature can be obtained by experimental testing; define the allowable working temperature of the missile wall cable as T a , and the initial temperature is generally selected as the ambient temperature T 0.

[0080] To establish the relevant first heat transfer finite element model, the density, specific heat capacity, and thermal conductivity of the insulation layer material are obtained; for the heat transfer process from the inner wall of the insulation layer to the cable on the missile wall, only heat conduction is considered. Through finite element calculation, for d in , within the time range of 0 to t act , when the maximum temperature on the side of the cable on the missile wall close to the inner wall of the insulation layer T inXL_Max does not exceed the allowable working temperature T a , d in the design theoretically meets the thermal protection requirements; to consider the design safety margin, the safety margin coefficient of the inner wall thickness of the insulation layer is defined as a in =|( T a - T inXL_Max ) / T inXL_Max |×100%; the larger this value is, d in the larger the

[0081] For the thickness from the outer wall of the insulation layer to the cable on the missile wall d out , the average heat flux density of the four regions on the outer wall of the engine housing is known, which are respectively q w_wind ( t ), q w_Rside ( t ), q w_lee ( t ), q w_ Lside ( t ); the allowable working temperature of the cable on the missile wall is T a , and the initial temperature is selected as the ambient temperature T 0.

[0082] To establish the relevant second heat transfer finite element model, in addition to the material parameters of the insulation layer, the material density, specific heat capacity, and thermal conductivity of the engine housing also need to be obtained; for the heat transfer process from the outer wall of the engine housing to the outer wall of the insulation layer, and then from the outer wall of the insulation layer to the cable on the missile wall, only heat conduction is considered; through finite element calculation, for d out , within the time range of 0 to t allWithin a certain time range, when the maximum temperature on the side of the missile wall cable close to the outer wall of the thermal insulation layer T outXL_Max does not exceed the allowable working temperature T a , d out the design theoretically meets the thermal protection requirements; to consider the design safety margin, define the safety margin coefficient of the inner wall thickness of the thermal insulation layer as a out =|( T a - T outXL_Max ) / T outXL_Max |×100%; the larger this value is, d out the larger the value is. Therefore, the selection of the safety margin should comprehensively consider the propellant charge of the engine and the outer wall thickness of the thermal insulation layer.

[0083] Step 3, locally set a thermal protection coating on the outer wall of the engine casing to further enhance the thermal protection ability of the thermal protection structure throughout the flight of the aircraft.

[0084] The circumferential distribution range requirement of the thermal protection coating on the outer wall of the engine casing is not only to completely cover the circumferential distribution range of the missile wall cable on the inner wall of the engine casing, but also to extend circumferentially at the circumferential boundary position of the thermal protection coating to improve the thermal protection margin at the boundary of the missile wall cable. The specific description is as follows:

[0085] Define the thickness of the thermal protection coating as d tc , the circumferential distribution angle of the thermal protection coating on the outer wall of the engine casing is φ tc , define the boundary margin of the missile wall cable wiring position on the engine casing as l , where ; to ensure that the thermal protection coating can completely cover the circumferential arrangement position of the missile wall cable in the engine casing and leave an appropriate margin at the boundary, then for φ tc , there is:

[0086] .

[0087] For the selection of the thickness d tc of the thermal protection coating, it is similar to the selection analysis process of d out in Step 2, and is obtained by establishing a heat transfer finite element model; in addition to the material parameters of the thermal insulation layer and the engine casing, it is also necessary to clarify the density, specific heat capacity, and thermal conductivity of the thermal protection coating; the third heat transfer finite element model established is the same as that in Step 2d out The difference is that the average heat flux density of four regions on the outer wall of the engine housing acts on the outer surface of the thermal protection coating; the allowable working temperature of the embedded cable is T a , and the initial temperature of the system is selected as the ambient temperature T 0.

[0088] For the heat transfer process from the thermal protection coating to the outer wall of the engine housing, from the outer wall of the engine housing to the outer wall of the insulation layer, and then from the outer wall of the insulation layer to the missile wall cable, only heat conduction is considered; through finite element calculation, for the thickness of the thermal protection coating d tc , within the range of 0 to t all time, when the highest temperature on the side of the missile wall cable close to the thermal protection coating after setting the thermal protection coating T tcXL_Max does not exceed the allowable working temperature T a , d tc the design theoretically meets the thermal protection requirements; to consider the design safety margin, the safety margin coefficient of the inner wall thickness of the insulation layer is defined as a tc =|( T a - T tcXL_Max ) / T tcXL_Max |×100%; the larger this value is, d tc the larger the d tc value is. However, in order to maintain the aerodynamic shape of the outer wall of the engine housing,

[0089] the d in , d in value cannot be too large.

[0090] In an embodiment of the present invention, referring to Figure 1 and Figure 2 , it is a design scheme for the thermal protection structure of the missile wall cable of the engine housing of a slender aircraft; to highlight the detailed features of the thermal protection structure, Figure 2 the dimensions of some components are enlarged in

[0091] First, the material of the engine casing 1 is steel. An insulation layer 2 is arranged between the engine charge 3 and the inner wall of the engine casing 1. The missile wall cable 4 is buried in the insulation layer 2 and runs through the engine casing 1 circumferentially, so that the front and rear ends of the missile wall cable 4 can be connected to the cable interfaces at the front and rear ends of the engine compartment section. Among them, the total thickness of the insulation layer 2 is 5 mm, the thickness from the inner wall of the insulation layer 2 to the missile wall cable 4 d in is 2 mm, and the thickness from the outer wall of the insulation layer 2 to the missile wall cable 4 d out is 3 mm. The reason for choosing ethylene propylene diene monomer (EPDM) rubber as the material of the insulation layer 2 is mainly due to its characteristics such as low density, high thermal decomposition temperature, large heat absorption value, good aging resistance and excellent mechanical properties. This material has been widely used in the insulation layer of solid aircraft engines.

[0092] Secondly, in the circumferential direction, looking forward from the rear of the engine casing 1, according to the principle of avoiding the overheat ablation area of the inner wall of the engine during the active flight section of the aircraft and selecting the area corresponding to the minimum average heat flux density on the inner wall of the engine casing 1, the missile wall cable 4 is arranged at a position where the angle between its center line and the OZ axis is 15°. This position is located on the right side of the four regions of the engine casing 1. The reason for choosing this region instead of the leeward side of the engine casing 1 is mainly determined by factors such as the change of the overheat ablation area of the engine and the complex pitch attitude change of the missile body during flight overload, which leads to the change of the heat flux density on the surface of the engine casing 1.

[0093] Finally, a 1-mm-thick TR48 thermal protection coating 5 is sprayed on the outer wall of the engine casing 1 corresponding to the center position of the missile wall cable 4. The angles of the left and right boundaries of the missile wall cable 4 from the center position of the cable are 20°. Considering the width of the distribution of the missile wall cable 4, the distribution range of the thermal protection coating 5 is: taking the line connecting the center position of the missile wall cable 4 and the center of the engine casing 1 as the reference, 30° to both sides, to ensure that the thermal protection coating 5 completely covers the position of the missile wall cable 4 circumferentially and improve the thermal protection ability of the missile wall cable 4. The thermal protection coating 5 is also a thermal protection coating with good performance and is widely used in the field of aircraft thermal protection.

[0094] Through finite element simulation analysis, during flight, the maximum temperature of the thermal protection coating 5 does not exceed 450 °C, while the allowable working temperature of the thermal protection coating 5 is 540 °C, and the safety margin can be more than 20%. By optimizing the thermal protection structure through simulation analysis, a large amount of design iteration time is saved. At the same time, through flight test verification, the missile wall cable 4 can work normally during the test, indicating that the thermal protection structure can meet the use requirements.

[0095] This structural solution avoids the changes in the aerodynamic shape caused by adding a cable cover to the outer wall of the engine casing 1, as well as the increase in the connection structure and processing cost. It is easy to process and assemble, has good processability, good adaptability to the launch box, and does not require modification of the interface of the launch box. In addition, the precise arrangement of the position of the thermal protection coating 5 can effectively improve the thermal protection ability of the cable, and at the same time, the influence on the aerodynamic shape and structural weight of the engine casing 1 is almost negligible.

[0096] The above embodiments are only used to illustrate the technical solutions of the present application, rather than to limit them; although the present application has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that: they can still modify the technical solutions recorded in the foregoing embodiments, or perform equivalent replacements on some of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present application, and should all be included in the protection scope of the present application.

Claims

1. A forward design method for the thermal protection structure of an engine casing elastic wall cable, characterized in that: include: The distribution of the overload ablation area of ​​the inner wall of the engine during the active flight phase of the aircraft and the distribution of the aerodynamic heating of the outer wall of the engine during the entire flight process of the aircraft are used as design constraints for the elastic wall cable to determine the layout position of the elastic wall cable in the circumferential direction of the engine casing; The elastic wall cable is arranged on the inner wall of the engine casing, and the elastic wall cable is wrapped in a heat protection material insulation layer to form a heat protection structure; wherein the outer wall of the insulation layer is in contact with the inner wall of the engine casing, and the inner wall of the insulation layer is in contact with the outer wall of the engine charge; The specific principles for determining the arrangement position of the elastic wall cable in the circumferential direction of the engine casing are as follows: Avoid the overload ablation area of ​​the engine inner wall and the affected adjacent areas during the active flight phase of the aircraft: in, Indicates existence, represents the layout angle of the elastic wall cable in the circumferential direction of the engine casing, k represents the boundary margin of the overload ablation area of ​​the engine inner wall, α represents the distribution angle of the overload ablation area of ​​the engine inner wall in the circumferential direction of the engine casing, and θ yz_Min ,θ yz_Max Indicates the minimum and maximum azimuth angles of the active segment of the aircraft; Among the four areas of the outer wall of the engine casing divided by counterclockwise rotation of the Y-axis and Z-axis of the missile body coordinate system: the windward side, the right side, the leeward side, and the left side, the area with the smallest average heat flux density is selected for layout; The circumferential distribution of the elastic cable on the inner wall of the engine casing should not interfere with other parts in the engine; The engine inner wall overload ablation area of ​​the flight active section of the aircraft and the affected adjacent area are specifically determined by: obtaining the circumferential position distribution of the engine inner wall overload ablation area on the inner wall of the engine casing according to the magnitude and direction of the flight lateral overload of the flight active section of the aircraft; Combined with the flight trajectory data of the aircraft, the trajectory is solved according to the flight mechanics principle of the aircraft. The flight lateral overload can be obtained by synthesizing the Y-direction overload and the Z-direction overload of the missile body coordinate system. Take the flight lateral overload azimuth angle θ yz The position of the inner wall of the engine casing corresponding to the opposite side is used as the overload ablation zone of the inner wall of the engine; The maximum flight time of the active phase of the aircraft is recorded as t act , when the flight lateral overload n yz As time t changes, the azimuth angle θ yz The range of change over time t is θ yz ∈[θ yz_Min ,θ yz_Max ], where t∈[0,t act ];θ yz_Min ,θ yz_Max Indicates the minimum and maximum azimuth angles of the active segment of the aircraft; Define the angle of the overload ablation area of ​​the engine inner wall in the circumferential direction of the engine casing as α, then: α=|θ yz_Max -θ yz_Min |; Define the boundary margin of the overload ablation area of ​​the engine inner wall as k, where k∈[0,1]; The overload ablation area of ​​the engine inner wall and the affected adjacent area S GZ Defined as: S GZ =[-kɑ+θ yz_Min ,kɑ+θ yz_Max ].

2. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 1 is characterized in that: The method for determining the four areas of the outer wall of the engine housing is as follows: Based on the aircraft body coordinate system, the Y-axis and Z-axis of the body coordinate system are rotated 45° counterclockwise in the YOZ plane. Then, the outer wall of the engine casing in the YOZ plane will be divided into four areas with a central angle of 90° by the rotated Y-axis and Z-axis, of which the area directly below is the windward side, the area on the right is the right side, the area directly above is the leeward side, and the area on the left is the left side.

3. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 1 is characterized in that: The method for determining the average heat flux is: The aerodynamic heating of the outer wall of the engine during the whole flight process of the aircraft is also combined with the flight trajectory data, and the surface heat flux density of the outer wall of the engine casing is obtained through the computational fluid dynamics method to obtain the change law of time; Divide the variation law into four zones according to the divided regions to obtain the heat flux density of each zone; The heat flux density in each area is then spatially averaged to obtain the average heat flux density of each area changing with time.

4. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 1 is characterized in that: In the protective structure, the thickness d from the inner wall of the insulation layer to the elastic wall cable in The method for determining is: Select the temperature T of the area close to the inner wall of the insulation layer in the engine casing in (t) is the temperature boundary condition of the inner wall area of ​​the insulation layer, where time t∈[0,t act ], t act is the maximum flight time of the active segment of the aircraft, and the allowable operating temperature of the elastic wall cable is defined as T a , the initial temperature is selected as the ambient temperature T0; The first heat transfer finite element model is established to obtain the density, specific heat capacity, and thermal conductivity of the insulation layer material. The heat transfer process from the inner wall of the insulation layer to the elastic wall cable only considers heat conduction. Through finite element calculation, for d in , in 0~t act Within the time range, when the elastic cable is close to the inner wall of the insulation layer, the highest temperature T inXL_Max Not exceeding the allowable operating temperature T a When in The design meets the thermal protection requirements.

5. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 1, characterized in that: In the protective structure, the thickness d from the outer wall of the insulation layer to the elastic wall cable out The method for determining is: The average heat flux density, insulation material parameters, material density, specific heat capacity, and thermal conductivity of the four regions of the engine casing outer wall are obtained; a second heat transfer finite element model is established, and the heat transfer process from the outer wall of the engine casing to the outer wall of the insulation layer and then from the outer wall of the insulation layer to the elastic wall cable also only considers heat conduction; through finite element calculation, for d out , in 0~t all Within the time range, when the maximum temperature T of the elastic cable close to the outer wall of the insulation layer outXL_Max Not exceeding the allowable operating temperature T a When out The design meets the thermal protection requirements; all Indicates the maximum time of the entire flight process of the aircraft.

6. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 1, characterized in that: The method also includes: locally setting a thermal protection coating on the outer wall of the engine casing; the circumferential distribution range of the thermal protection coating on the outer wall of the engine casing not only needs to completely cover the circumferential distribution range of the elastic wall cable on the inner wall of the engine casing, but also extends circumferentially at the circumferential boundary position of the thermal protection coating.

7. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 6 is characterized in that: The angle of the circumferential distribution of the thermal protection coating on the outer wall of the engine casing is defined as Define the boundary margin of the cable routing position of the engine casing elastic wall as l, where l∈[0,1]; then: Where β represents the arrangement angle of the elastic wall cable in the circumferential direction of the engine casing. They respectively represent the maximum and minimum values ​​of the circumferential arrangement angle of the elastic wall cable in the engine casing.

8. The forward design method for the thermal protection structure of the elastic wall cable of the engine casing according to claim 6 is characterized in that: For the thickness of the thermal protection coating d tc The selection of is obtained by establishing the third heat transfer finite element model; When modeling, in addition to the material parameters of the insulation layer and the engine casing, it is also necessary to clarify the density, specific heat capacity, and thermal conductivity of the thermal protection coating; the average heat flux density of the four regions of the outer wall of the engine casing is applied to the outer surface of the thermal protection coating; the heat transfer process from the thermal protection coating to the outer wall of the engine casing, from the outer wall of the engine casing to the outer wall of the insulation layer, and then from the outer wall of the insulation layer to the elastic wall cable only considers heat conduction; Through finite element calculation, the thickness d of the thermal protection coating tc , in 0~t all Within the time range, when the thermal protection coating is set, the maximum temperature T of the elastic wall cable close to the thermal protection coating side tcXL_Max Not exceeding the allowable operating temperature T a When tc The design meets the thermal protection requirements; all Indicates the maximum time of the entire flight process of the aircraft.

Images