Forward design method for thermal protection structure of engine shell elastic wall cable

Abstract

The invention discloses a thermal protection structure forward design method for an engine shell elastic wall cable, which comprises the following steps: taking the distribution of an engine inner wall overload ablation area of an aircraft flight active section and the distribution of engine outer wall pneumatic heating in the whole aircraft flight process as design constraints of the elastic wall cable; the arrangement position of the elastic wall cable in the circumferential direction of the engine shell is determined; the mode that an elastic wall cable is arranged on the inner wall of an engine shell is adopted, and the elastic wall cable is wrapped in a thermal protection material heat insulation layer to form a thermal protection structure; wherein the outer wall of the heat insulation layer is attached to the inner wall of the engine shell, and the inner wall of the heat insulation layer is attached to the outer wall of the charging position of the engine. According to the method, the problems that the number of iteration times of the scheme is large, the design period is long and the design cost is increased due to the fact that traditional engineering experience is dominated and repeated iterative design is adopted are effectively solved.

Description

Technical Field

[0001] The present invention relates to the field of thermal protection of aircraft engines, and in particular to a forward design method for a thermal protection structure of an engine casing elastic wall cable. Background Art

[0002] For aircraft, there are currently two main types of engine casing elastic wall cables commonly used. One type is to arrange the elastic wall cables on the outside of the engine casing, and use a cable heat insulation cover on the outer wall of the engine casing to protect the cables on the outer wall of the engine casing; the other type is to arrange the elastic wall cables on the inside of the engine casing, and use an insulation layer to protect the inner wall cables.

[0003] Regarding the solution of arranging the elastic wall cables on the outside of the engine casing and using a cable heat shield to protect the cables, since the heat shield forms a local bulge in the circumference of the engine casing, during the flight of the aircraft, the aerodynamic heating at the bulge position of the heat shield is enhanced, the local electromagnetic environment is deteriorated, and the aerodynamic shearing effect at the connection between the heat shield and the outer wall of the engine casing is also strong. The following problems exist: In terms of the aerodynamic shape of the aircraft, the use of the outer wall cable heat shield structure will cause a bulge in the circumference of the outer wall of the engine casing, which is not conducive to maintaining the overall aerodynamic shape of the aircraft engine compartment section, which may have an impact on the aerodynamic performance of the aircraft. Bring adverse effects; in terms of structural design and manufacturing process, the cable heat insulation cover must have strong thermal protection capabilities and anti-electromagnetic interference capabilities, and the connection between the engine and the engine casing must be strengthened; in terms of adaptability to the launch platform, the structural scheme of the outer wall cable heat insulation cover is adopted, and the local bulge caused by it may be difficult to match the interface size of the launch platform; in terms of storage and transportation, the raised outer wall cable heat insulation cover is at risk of bumping during storage and transportation, and additional buffering protection treatment is required, which has higher requirements on the impact resistance of the storage and transportation box.

[0004] For the scheme of arranging the elastic wall cable inside the engine casing and using the insulation layer to protect the inner wall cable, this kind of scheme has high requirements on the insulation effect of the insulation layer and the specific arrangement position of the cable on the circumference of the inner wall of the engine casing. In particular, in terms of the selection of the arrangement orientation of the cable on the inner wall of the engine casing, how to choose the appropriate 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, there are currently no clear design specifications and specific requirements or guidance. Therefore, in actual engineering applications, it is often necessary to rely on engineering experience for design, which may result in a large number of scheme iterations and repeated designs, resulting in a long design cycle and increased design costs, which 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 an engine casing elastic wall cable, so as to overcome the problems existing in the wiring process of the existing engine casing elastic wall cable.

[0006] In order to achieve the above tasks, the present invention adopts the following technical solutions: A forward design method for a thermal protection structure of an engine casing elastic wall cable, comprising: 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 wrapped in an insulating layer of thermal protection material to form a thermal protection structure; wherein the outer wall of the insulating layer is in contact with the inner wall of the engine casing, and the inner wall of the insulating layer is in contact with the outer wall of the engine charge.

[0007] Furthermore, 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, Indicates the layout angle of the elastic wall cable in the circumference of the engine casing, k Indicates the boundary margin of the overload ablation area on the inner wall of the engine, α Indicates the angle of the overload ablation area on the inner wall of the engine distributed around the engine casing. , Indicates the minimum and maximum azimuth angles of the active segment of the aircraft; According to the missile body coordinate system Y Axis and Z Among the four areas of the outer wall of the engine casing divided by the counterclockwise rotation of the axis: 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 position distribution of the elastic wall cable on the inner wall of the engine casing should not interfere with other components in the engine.

[0008] Furthermore, the overload ablation area of ​​the engine inner wall of the active flight section of the aircraft and the affected adjacent area are specifically determined by: According to the magnitude and direction of the flight lateral overload of the aircraft during the active flight phase, the circumferential position distribution of the overload ablation area of ​​the engine inner wall on the inner wall of the engine casing is obtained; Combined with the aircraft flight trajectory data, the trajectory is solved according to the principles of aircraft flight mechanics. Y To overload and Z By combining the lateral overload, we can get the lateral overload in flight. Take the flight lateral overload azimuth 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 segment of the aircraft is recorded as , when the flight lateral overload Over time t When the azimuth changes Also over time t The range of variation is ,in ; Indicates the minimum and maximum azimuth angles of the active segment of the aircraft; The angle of the overload ablation area on the inner wall of the engine in the circumferential direction of the engine casing is defined as α , then: ; The boundary margin of the overload ablation area on the inner wall of the engine is defined as k ,in ; Overload ablation area of ​​the engine inner wall and the affected adjacent areas S GZ Defined as: .

[0009] Furthermore, the method for determining the four areas of the outer wall of the engine casing is: Based on the aircraft body coordinate system, YOZ The coordinate system of the projectile body is Y Axis and Z The axis rotates 45° counterclockwise. YOZ The outer wall of the engine housing in the plane will be rotated Y Axis and Z The axis is divided into four areas with a central angle of 90°, 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.

[0010] Furthermore, the average heat flux density is determined by: 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; The variation law is divided into four zones according to the divided regions to obtain the heat flux density of each zone; the heat flux density of each zone is spatially averaged to obtain the average heat flux density of each zone varying with time.

[0011] Furthermore, in the protective structure, the thickness from the inner wall of the insulation layer to the elastic wall cable is din The method to determine is: Select the temperature of the area inside the engine casing close to the inner wall of the insulation layer T in ( t ) as the temperature boundary condition of the inner wall area of ​​the insulation layer, where time , is the maximum flight time of the active segment of the aircraft, and the allowable operating temperature of the elastic cable is defined as T a , the initial temperature is selected as the ambient temperature T 0; 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, d in , at 0~ t act Within the time range, the maximum temperature of the elastic cable close to the inner wall of the insulation layer T inXL_Max Do not exceed the allowable operating temperature T a hour, d in The design meets the thermal protection requirements.

[0012] Furthermore, in the protective structure, the thickness from the outer wall of the insulation layer to the elastic wall cable is d out The method to determine 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 , at 0~ t all Within the time range, the maximum temperature of the elastic cable close to the outer wall of the insulation layer T outXL_Max Do not exceed the allowable operating temperature T a hour, d out The design meets the thermal protection requirements; t all Indicates the maximum time of the entire flight process of the aircraft.

[0013] Furthermore, 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.

[0014] Furthermore, the circumferential distribution angle of the thermal protection coating on the outer wall of the engine casing is defined as φ tc , define the margin of the cable routing position of the engine casing elastic wall as l ,in ; then: ; in, β Indicates the layout angle of the elastic wall cable in the circumference 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.

[0015] Furthermore, 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 areas 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 of the thermal protection coating d tc , at 0~ t all The maximum temperature of the side of the elastic cable close to the thermal protection coating after the thermal protection coating is set within the time range. T tcXL_Max Do not exceed the allowable operating temperature T a hour, d tc The design meets the thermal protection requirements; t all Indicates the maximum time of the entire flight process of the aircraft.

[0016] Compared with the prior art, the present invention has the following technical features: 1. Compared with the structural solution of arranging the elastic wall cables on the outside of the engine casing and using a heat insulation cover on the outer wall of the engine casing to protect the outer wall cables, the design of the present invention has the characteristics of being able to maintain the aerodynamic shape of the aircraft engine casing, light structure weight, good processability, low cost, easy installation, strong adaptability to the launch platform, and good storage and transportation performance.

[0017] 2. Compared with the solution of arranging the elastic wall cable on the inner side of the engine casing and using an insulating layer to protect the inner wall cable, the present invention also uses a thinner thermal protection coating material on the local outer wall of the engine casing corresponding to the position of the elastic wall cable, further enhancing the thermal protection capability of the elastic wall cable during the entire flight process of the aircraft.

[0018] 3. The structural thermal protection forward design method of the present invention establishes a finite element simulation model, takes the thermal environment of the inner and outer walls of the engine housing as the design input, clarifies the boundary conditions of the cable thermal protection structure, calculates and analyzes the relevant parameters of the structure, optimizes 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 problem of the traditional repeated iterative design dominated by engineering experience, resulting in a large number of solution iterations, a long design cycle, and increased design costs. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 is the aircraft body coordinate system YOZ Schematic diagram of the four area division in the plane.

[0020] Figure 2 Schematic diagram of a thermal protection structure in one embodiment of the present invention.

[0021] Explanation of reference numerals: 1 engine casing, 2 insulation layer, 3 engine charge, 4 bomb wall cable, 5 thermal protection coating. DETAILED DESCRIPTION

[0022] Based on the scheme of arranging elastic-wall cables on the inner side of an engine casing and using an insulating layer to protect the inner wall cables, and on the basis of the advantages of being able to maintain the aerodynamic shape of an aircraft engine casing, light structural weight, good processability, low cost, and easy installation, the present invention provides a forward design method for elastic-wall cables of an engine casing. Based on the idea of ​​forward design, the thermal environment of the inner and outer walls of the engine casing during the flight of the aircraft is used as a design input, and simulation analysis is performed by establishing a finite element model to obtain relevant parameters of the cable thermal protection structure, and the thermal protection structure is optimized and designed, thereby solving the problems of a large number of solution iterations, a long design cycle, and increased design costs caused by traditional iterative designs dominated by engineering experience.

[0023] The present invention provides a forward design method for a thermal protection structure of an engine casing elastic wall cable, comprising the following steps: Step 1: Taking the distribution of the overload ablation area of ​​the engine inner wall during the active flight phase of the aircraft and the distribution of the aerodynamic heating of the engine outer wall during the entire flight process of the aircraft as design constraints for the elastic wall cable, determine the layout position of the elastic wall cable in the circumferential direction of the engine casing.

[0024] Generally speaking, the distribution of the overload ablation area on the inner wall of the engine during the active flight phase of the aircraft and the most serious windward area of ​​the aerodynamic heating of the outer wall of the engine during the entire flight process of the aircraft are different in the circumferential position of the engine casing. Therefore, the circumferential arrangement position of the elastic wall cable should avoid the above two high-temperature areas as much as possible.

[0025] (1.1) The overload ablation area on the inner wall of the engine during the active flight phase of the aircraft.

[0026] According to the size and direction of the flight lateral overload in the active phase of the flight of the aircraft, the circumferential position distribution of the overload ablation area of ​​the inner wall of the engine can be obtained; the position of the overload ablation area of ​​the inner wall of the engine is generally on the opposite side of the flight lateral overload. The flight lateral overload can be combined with the flight trajectory data of the aircraft, and the trajectory can be solved according to the flight mechanics principle of the aircraft. Y Overload n y and Z Overload n z Synthesis, can get flight lateral overload n yz Size: ; The body coordinate system is based on the axial direction of the aircraft. X Axis, the cross section of the aircraft is based on the position of the aircraft axis O Point, build Y Axis and Z Axis, such as Figure 1 shown.

[0027] Take flight lateral overload n yz Azimuth The position of the inner wall of the engine casing corresponding to the opposite side is taken as the overload ablation zone of the inner wall of the engine; wherein the azimuth angle It is expressed as follows: ; The maximum flight time of the active segment of the aircraft is recorded as t act The maximum time of the entire flight process is t all , when the flight lateral overload Over time tWhen the azimuth changes Also over time t Changes, the range of which is , which is the circumferential distribution of the overload ablation area on the inner wall of the engine, where ; Indicates the minimum and maximum azimuth angles of the active segment of the aircraft.

[0028] The angle of the overload ablation area on the inner wall of the engine is defined as α , then: ; The boundary margin of the overload ablation area on the inner wall of the engine is defined as k ,in ; Considering the impact of the boundary of the overload ablation area on the inner wall of the engine on the adjacent area, in order to enhance the safety margin of the design, the layout area of ​​the elastic wall cable should also avoid the adjacent area; The overload ablation area of ​​the inner wall of the engine and the affected adjacent area S GZ Defined as: ; For example, when k When it is 0.2, it means that the influence range of the overload ablation area of ​​the inner wall of the engine on the adjacent area of ​​the single-side boundary accounts for 20% of the circumferential distribution angle range of the ablation overload area, and the interval that the circumferential arrangement position of the elastic wall cable needs to avoid is S GZ ; k The larger the value is, the less likely the circumferential arrangement of the elastic wall cable will be affected by the engine overload ablation zone, and the greater the safety margin of the design.

[0029] (1.2) Aerodynamic heating of the engine outer wall during the entire flight process of the aircraft.

[0030] The aerodynamic heating of the outer wall of the engine during the entire 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 by the computational fluid dynamics (CFD) method. q w Over time t The changing pattern of q w ( t ).

[0031] For ease of understanding, based on the aircraft body coordinate system, YOZ The coordinate system of the projectile body is Y Axis and Z The axis rotates 45° counterclockwise. YOZ The outer wall of the engine housing in the plane will be rotated Y Axis and ZThe axis is divided into four areas with a central angle of 90°, of which the windward side is directly below S wind , the right area is the right side S Rside , the area directly above is the leeward side S lee , the left area is the left side S Lside ,like Figure 1 shown.

[0032] For the convenience of calculation, the surface heat flux density of the outer wall of the engine casing is q w Over time t The law of change q w ( t ) According to the above four regions, the heat flux density of each region is obtained; then the heat flux density of each region is spatially averaged, and the heat flux density of each region over time is obtained. t The average heat flux density of the windward side q w_wind ( t ), the average heat flux density on the right side q w_Rside ( t ), the average heat flux density on the leeward side q w_lee ( t ), the average heat flux density on the left side q w_ Lside ( t ).

[0033] For most situations during the flight of an aircraft, the law of average heat flux density is generally: q w_wind ( t )> q w_Rside ( t )≈ q w_ Lside ( t )> q w_lee ( t ); however, in practical applications, a comprehensive analysis should be conducted in combination with ballistic flight data. For example, when the projectile undergoes axial rolling, pitching, and yaw motion during flight, the above rules may no longer apply; the circumferential arrangement position of the cable on the wall should be selected from the area with the smallest surface heat flux density among the above four areas.

[0034] When designing the circumferential arrangement position of the elastic wall cable, in order to improve the thermal protection capability of the cable as a whole in the radial direction of the engine casing, the elastic wall cable is arranged flatly along the circumferential position; the maximum and minimum values ​​of the circumferential arrangement angle of the elastic wall cable in the engine casing are defined as , , then the angle of the circumferential centerline of the elastic cable is ,in, =1 / 2( + ); define the range of the circumferential distribution angle of the elastic wall cable as S XL =[ , ], the layout angle of the elastic wall cable around the engine casing , both have .

[0035] The arrangement angle of the elastic wall cable in the circumference of the engine casing is defined as β , then: .

[0036] In summary, the design principles for the circumferential arrangement of the elastic wall cable in the engine casing are: (1) Avoid the overload ablation area on the inner wall of the engine during the active flight phase of the aircraft and the affected adjacent areas, namely: .

[0037] (2) According to the projectile coordinate system Y Axis and Z Among the four areas of the outer wall of the engine casing divided by counterclockwise rotation of the axis: 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.

[0038] (3) Considering the installation convenience of the elastic wall cable, its circumferential position distribution on the inner wall of the engine casing should not interfere with other components in the engine.

[0039] Step 2, arrange the elastic wall cable on the inner wall of the engine casing, and wrap the elastic wall cable in a thermal protection material insulation layer to form a thermal protection structure to maintain the circumferential consistency of the aerodynamic shape of the engine casing; 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.

[0040] The thickness from the inner wall of the insulation layer to the elastic cable d in, is to take the temperature of the engine charge at the active stage of the aircraft flight as the calculation input, and the temperature of the elastic wall cable does not exceed the allowable operating temperature as the constraint design; the thickness of the outer wall of the insulation layer to the elastic wall cable d out , is obtained by taking the heat flux density of the outer wall of the engine casing as the calculation input during the entire flight of the aircraft, and the temperature of the elastic wall cable not exceeding the allowable operating temperature as the constraint design. The specific description is as follows: For the thickness of the inner wall of the insulation layer to the elastic wall cable d in The engine charge continues to burn during the active phase of the flight. To simplify the calculation model, the temperature of the area close to the inner wall of the insulation layer in the engine casing is selected. T in ( t ) is used as the temperature boundary condition of the inner wall area of ​​the insulation layer, where ; This temperature can be obtained by testing; the allowable operating temperature of the elastic wall cable is defined as T a The initial temperature is usually selected as the ambient temperature T 0.

[0041] In order to establish the first heat transfer finite element model, the density, specific heat capacity and thermal conductivity of the insulation layer material are obtained; the heat transfer process from the inner wall of the insulation layer to the elastic wall cable only considers heat conduction, and through finite element calculation, d in , at 0~ t act Within the time range, the maximum temperature of the elastic cable close to the inner wall of the insulation layer T inXL_Max Do not exceed the allowable operating temperature T a hour, d in The design meets the thermal protection requirements in theory. 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 the value, d in Therefore, the selection of the full margin coefficient should comprehensively consider the engine charge and the thickness of the inner wall of the insulation layer.

[0042] For the thickness of the outer wall of the insulation layer to the elastic wall cable d outThe average heat flux density of the four regions of the engine casing outer wall is known, which are q w_wind ( t ), q w_Rside ( t ), q w_lee ( t ), q w_ Lside ( t ); The allowable operating temperature of the elastic wall cable is T a , the initial temperature is selected as the ambient temperature T 0.

[0043] In order to establish the relevant second heat transfer finite element model, in addition to the insulation material parameters, it is also necessary to obtain the material density, specific heat capacity, and thermal conductivity of the engine casing; 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 , at 0~ t all Within the time range, the maximum temperature of the elastic cable close to the outer wall of the insulation layer T outXL_Max Do not exceed the allowable operating temperature T a hour, d out The design meets the thermal protection requirements in theory. To consider the design safety margin, the safety margin coefficient of the inner wall thickness of the insulation layer is defined as a out =|( T a - T outXL_Max ) / T outXL_Max |×100%; the larger the value, d out Therefore, the selection of safety margin should take into account the engine charge and the thickness of the outer wall of the insulation layer.

[0044] Step 3: A thermal protection coating is partially provided on the outer wall of the engine casing, thereby further enhancing the thermal protection capability of the thermal protection structure during the entire flight process of the aircraft.

[0045] The circumferential distribution range of the thermal protection coating on the outer wall of the engine casing is required to not only completely cover the circumferential distribution range of the elastic wall cable on the inner wall of the engine casing, but also extend circumferentially at the circumferential boundary position of the thermal protection coating to increase the thermal protection margin at the boundary of the elastic wall cable. The specific description is as follows: The thickness of the thermal protective coating is defined as d tc The angle of the circumferential distribution of the thermal protection coating on the outer wall of the engine casing is φ tc , define the margin of the cable routing position of the engine casing elastic wall as l ,in To ensure that the thermal protection coating can completely cover the circumferential arrangement position of the elastic wall cable in the engine casing and leave a suitable margin at the boundary, φ tc ,have: .

[0046] For the thickness of thermal protective coating d tc The selection is the same as in step 2 d out The selection and analysis process is similar, and both are obtained by establishing a heat transfer finite element model; 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 third heat transfer finite element model established is the same as that in step 2 d out The difference is that the average heat flux density of the four areas of the outer wall of the engine casing acts on the outer surface of the thermal protection coating; the allowable operating temperature of the buried cable is T a , the initial temperature of the system is selected as the ambient temperature T 0.

[0047] 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 also only considers heat conduction; through finite element calculation, the thickness of the thermal protection coating d tc , at 0~ t all The maximum temperature of the side of the elastic cable close to the thermal protection coating after the thermal protection coating is set within the time range. T tcXL_Max Do not exceed the allowable operating temperature T a hour, d tc The design meets the thermal protection requirements in theory. 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 the value,d tc The larger the value, however, in order to maintain the aerodynamic shape of the outer wall of the engine casing, d tc The value cannot be too large.

[0048] In the insulation layer d in , d in When all conditions are determined, adding a thermal protection coating of appropriate thickness can help improve the thermal protection capability of elastic wall cables.

[0049] In one embodiment of the present invention, see Figure 1 and Figure 2 , is a design scheme for the thermal protection structure of the engine case, wall and cable of a slender aircraft; in order to highlight the detailed features of the thermal protection structure, Figure 2 The dimensions of some components are enlarged; in this embodiment, the heat insulation layer 2 is made of EPDM rubber, and the heat protection coating is made of TR48 heat protection material. The scheme of the heat protection structure is as follows: First, the material of the engine casing 1 is steel, and an insulation layer 2 is arranged between the engine charge 3 and the inner wall of the engine casing 1. The elastic wall cable 4 is buried in the insulation layer 2 and penetrates the engine casing 1 circumferentially, so that the front and rear ends of the elastic wall cable 4 can be connected to the front and rear end cable interfaces of the engine compartment; wherein, the total thickness of the insulation layer 2 is 5 mm, and the thickness from the inner wall of the insulation layer 2 to the elastic wall cable 4 is d in 2mm, the thickness from the outer wall of the insulation layer 2 to the elastic wall cable 4 d out The reason why EPDM rubber is used as the material of the insulation layer 2 is mainly due to its low density, high thermal decomposition temperature, large heat absorption value, aging resistance and excellent mechanical properties. This material has been widely used in the insulation layer of solid aircraft engines.

[0050] Secondly, in the circumferential direction, looking forward from the rear of the engine casing 1, according to the principle of avoiding the overload ablation area of ​​the inner wall of the engine in the active flight section of the aircraft and selecting the inner wall of the engine casing 1 corresponding to the area with the smallest average heat flux density, the elastic wall cable 4 is arranged at a distance of 1.5 m from its center line. OZ The position with an axis angle of 15° is located on the right side of the four areas of the engine casing 1; the reason for selecting this area instead of the leeward side of the engine casing 1 is mainly determined by factors such as changes in the engine overload ablation zone and complex pitch attitude changes of the projectile during flight overload, which lead to changes in the heat flux density on the surface of the engine casing 1.

[0051] 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 elastic wall cable 4. The left and right boundaries of the elastic wall cable 4 are 20° away from the center position of the cable. Combined with the width of the distribution of the elastic wall cable 4, the distribution range of the thermal protection coating 5 is: based on the line connecting the center position of the elastic wall cable 4 and the center of the engine casing 1, 30° on both sides to ensure that the thermal protection coating 5 completely covers the position of the elastic wall cable 4 in the circumferential direction, thereby improving the thermal protection capability of the elastic wall cable 4; the thermal protection coating 5 is also a thermal protection coating with better performance and is widely used in the field of aircraft thermal protection.

[0052] According to finite element simulation analysis, during the flight, the maximum temperature of the thermal protection coating 5 does not exceed 450°C, while the allowable operating temperature of the thermal protection coating 5 is 540°C, and the safety margin can be increased by more than 20%. The thermal protection structure is optimized through simulation analysis, which saves a lot of design iteration time. At the same time, after flight test verification, the elastic wall cable 4 can work normally in the test, indicating that the thermal protection structure can meet the use requirements.

[0053] This structural solution avoids the change in aerodynamic shape caused by adding a cable cover to the outer wall of the engine housing 1, as well as the increase in connection structure and processing costs. It is easy to process and assemble, has good processability, and has good adaptability to the launch box, and does not require changes to 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 capability of the cable, while the impact on the aerodynamic shape and structural weight of the engine housing 1 is almost negligible.

[0054] 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 aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. These modifications or replacements do not deviate the essence of the corresponding technical solutions 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 wrapped in an insulating layer of thermal protection material to form a thermal protection structure; wherein the outer wall of the insulating layer is in contact with the inner wall of the engine casing, and the inner wall of the insulating layer is in contact with the outer wall of the engine charge.

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 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, Indicates the layout angle of the elastic wall cable in the circumference of the engine casing, k Indicates the boundary margin of the overload ablation area on the inner wall of the engine, Indicates the angle of the overload ablation area on the inner wall of the engine distributed around the engine casing. , Indicates the minimum and maximum azimuth angles of the active segment of the aircraft; According to the missile body coordinate system Y Axis and Z The four areas of the outer wall of the engine casing divided by the counterclockwise rotation of the axis: the windward side, the right side, the leeward side, and the left side, select the area with the smallest average heat flux density for layout; The circumferential position distribution of the elastic wall cable on the inner wall of the engine casing should not interfere with other components in the engine.

3. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 2 is characterized in that: The overload ablation area of ​​the engine inner wall of the active flight section of the aircraft and the affected adjacent area are specifically determined by: According to the magnitude and direction of the flight lateral overload of the aircraft during the active flight phase, the circumferential position distribution of the overload ablation area of ​​the engine inner wall on the inner wall of the engine casing is obtained; Combined with the aircraft flight trajectory data, the trajectory is solved according to the principles of aircraft flight mechanics. Y To overload and Z By combining the lateral overload, we can get the lateral overload in flight. Take the flight lateral overload azimuth 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 segment of the aircraft is recorded as , when the flight lateral overload Over time t When the azimuth changes Also over time t The range of variation is ,in ; Indicates the minimum and maximum azimuth angles of the active segment of the aircraft; The angle of the overload ablation area on the inner wall of the engine in the circumferential direction of the engine casing is defined as , then: ; The boundary margin of the overload ablation area on the inner wall of the engine is defined as k ,in ; Overload ablation area of ​​the engine inner wall and the affected adjacent areas S GZ Defined as: .

4. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 2 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, YOZ The coordinate system of the projectile body is Y Axis and Z The axis rotates 45° counterclockwise. YOZ The outer wall of the engine housing in the plane will be rotated Y Axis and Z The axis is divided into four areas with a central angle of 90°, 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.

5. The forward design method for the thermal protection structure of the engine casing elastic wall cable according to claim 2 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.

6. The forward design method for the thermal protection structure of the elastic wall cable of the engine casing according to claim 1 is characterized in that: In the protective structure, the thickness of the inner wall of the insulation layer to the elastic wall cable The method to determine is: Select the temperature of the area inside the engine casing close to the inner wall of the insulation layer As the temperature boundary condition of the inner wall area of ​​the insulation layer, the time , is the maximum flight time of the active segment of the aircraft, and the allowable operating temperature of the elastic cable is defined as , the initial temperature is selected as the ambient temperature ; 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, , at 0~ Within the time range, the maximum temperature of the elastic cable close to the inner wall of the insulation layer Do not exceed the allowable operating temperature hour, The design meets the thermal protection requirements.

7. The forward design method for the thermal protection structure of the elastic wall cable of the engine casing according to claim 2 is characterized in that: In the protective structure, the thickness of the outer wall of the insulation layer to the elastic wall cable The method to determine is: The average heat flux density, insulation material parameters, material density, specific heat capacity, and thermal conductivity coefficient 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, , at 0~ Within the time range, the maximum temperature of the elastic cable close to the outer wall of the insulation layer Do not exceed the allowable operating temperature hour, The design meets the thermal protection requirements; Indicates the maximum time of the entire flight process of the aircraft.

8. The forward design method for the thermal protection structure of the elastic wall cable of the engine casing according to claim 1 is 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.

9. The forward design method for the thermal protection structure of the elastic wall cable of the engine casing according to claim 8 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 margin of the cable routing position of the engine casing elastic wall as l ,in ; then: ; in, Indicates the layout angle of the elastic wall cable in the circumference 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.

10. The forward design method for the thermal protection structure of the elastic wall cable of the engine casing according to claim 8, characterized in that: For the thickness of thermal protective coating 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 of the thermal protection coating , at 0~ The maximum temperature of the side of the elastic cable close to the thermal protection coating after the thermal protection coating is set within the time range. Do not exceed the allowable operating temperature hour, The design meets the thermal protection requirements; Indicates the maximum time of the entire flight process of the aircraft.

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