A morphing aircraft system model for wind tunnel testing

Abstract

The application discloses a deformable aircraft system model for a wind tunnel test, which comprises an aircraft main body structure model, an external supporting rod structure, a balance measurement system, a wing surface transmission system and an aircraft internal partition plate; wherein one end of the external supporting rod structure is connected with a tail end of the aircraft main body structure model, and the other end of the external supporting rod structure is connected with an angle of attack mechanism; the balance measurement system and the wing surface transmission system are arranged at the rear end of the inside of the aircraft main body structure model; the balance measurement system is located at the upper part of the wing surface transmission system, and the aircraft internal partition plate is arranged between the balance measurement system and the wing surface transmission system. The application can simultaneously accommodate the wing surface transmission system and the balance measurement system in a small aircraft, and solves the interference problem of the wing surface transmission system to the balance measurement system.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of hypersonic wind tunnel test, and particularly relates to a deformable aircraft system model for wind tunnel test. BACKGROUND

[0002] Traditional fixed configuration aircraft are usually designed with a fixed design point, and can only maintain high performance in a certain flight profile, but the flight performance will sharply decrease beyond the design point, and even may cause the risk of aircraft losing control. The deformable aircraft is a combination of different configuration aircraft aerodynamic shape characteristics, and actively changes the aerodynamic shape according to the change of flight environment, greatly improves the whole task flight ability, and realizes the best flight performance in different flow states. The development of deformable aircraft has become an important way to break through the performance of future aircraft, and has attracted the attention of many scholars at home and abroad.

[0003] The common aircraft deformation methods are folding wings and turning wings. On the one hand, the deformation principle of folding / turning wings is relatively simple, and on the other hand, the wing increasing effect is obvious. Extensive research has been carried out on the above two types of deformers. At present, the research of deformable aircraft at home and abroad mainly focuses on the simulation design of wing surface deformation mechanism, computer numerical simulation analysis and different configuration steady structure analysis. The common wing surface deformation mechanism includes single folding wing, Z-shaped folding wing, scissors type deformation wing, horizontal embedded deformation wing, etc. By using computer simulation analysis technology, it is found that the deformable aircraft can significantly improve the overall performance of the aircraft. At the same time, for the aircraft deformation process, the wing surface deformation direction, deformation stability, deflection analysis, wing surface motion time, mass layout and other key parameters are actively explored. In order to better evaluate the actual deformation effect of the deformable aircraft, the ground wind tunnel test is adopted to verify and analyze: the wind tunnel force test of the different configuration steady structure of the deformable aircraft model is carried out, the aerodynamic characteristics of the model under different attack angles, folding modes, Ma numbers and roll angles are studied, and the test results show that the wind tunnel test results are in good agreement with the simulation results, and the deformable aircraft has excellent aerodynamic performance.

[0004] However, the current research on the dynamic aerodynamic characteristics of the morphing aircraft during the morphing process is relatively insufficient. Due to the large-scale shape change of the aircraft, on the one hand, the aircraft cannot avoid the local structure such as the discontinuity of the shape surface and the connection gap, and on the other hand, the flow field state changes accordingly. The existence of the above phenomena leads to the complexity of the aerodynamic characteristics and the environmental load of the morphing aircraft. Therefore, the existing problems of the morphing aircraft are as follows: the ground wind tunnel test method cannot be used to measure the aerodynamic parameters and the large-scale shape change of the morphing aircraft at the same time; the temperature of the hypersonic wind tunnel is relatively high, and it is difficult to verify the morphing effect of the high Mach number aircraft by using the test method; the internal space of the morphing aircraft is limited, and it is difficult to simultaneously accommodate the wing transmission system and the balance measurement system, and it is difficult to avoid the internal interference between the two systems. SUMMARY

[0005] The technical problem solved by the present application is to overcome the shortcomings of the prior art and provide a morphing aircraft system model for wind tunnel test, which can simultaneously accommodate the wing transmission system and the balance measurement system in a small aircraft, solves the interference problem of the wing transmission system to the balance measurement system, and realizes the accurate measurement of the aerodynamic force of the morphing aircraft.

[0006] The object of the present application is achieved by the following technical scheme: a morphing aircraft system model for wind tunnel test, comprising: an aircraft main body structure model, an external connecting strut structure, a balance measurement system, a wing transmission system and an aircraft internal partition; wherein one end of the external connecting strut structure is connected with the tail end of the aircraft main body structure model, and the other end of the external connecting strut structure is connected with an angle of attack mechanism; the balance measurement system and the wing transmission system are arranged at the rear end of the inside of the aircraft main body structure model; the balance measurement system is located above the wing transmission system, and an aircraft internal partition is arranged between the balance measurement system and the wing transmission system.

[0007] In the above morphing aircraft system model for wind tunnel test, the scale of the morphing aircraft system model is obtained according to the blockage of the model.

[0008] In the above morphing aircraft system model for wind tunnel test, the scale of the morphing aircraft system model is obtained by the following formula:

[0009] ε=k[1-q(λ) / σ];

[0010] q(λ)=A M / A T ;

[0011] Wherein, A Mis the first throat cross-sectional area of the wind tunnel; q(λ) is a gas dynamic function; λ is the wavelength of the gas; σ is the total pressure recovery coefficient of the normal supersonic shock; and k is a constant.

[0012] In the deformable aircraft system model for wind tunnel test, the constant k is 0.25-0.40.

[0013] In the deformable aircraft system model for wind tunnel test, the size of the deformable aircraft system model is:

[0014] V t = V·ε

[0015] In the formula, V t is the size of the deformable aircraft system model, V is the volume of the aircraft, and ε is the scale of the deformable aircraft system model.

[0016] In the deformable aircraft system model for wind tunnel test, the internal space of the deformable aircraft system model is obtained according to the size of the deformable aircraft system model, wherein the internal space of the deformable aircraft system model is calculated according to the following formula:

[0017] V i = V t ·n

[0018] In the formula, V i is the internal space of the deformable aircraft system model, and n is a constant.

[0019] In the deformable aircraft system model for wind tunnel test, the internal space of the balance measurement system and the internal space of the wing transmission system are obtained according to the internal space of the deformable aircraft system model, wherein the internal space of the balance measurement system and the internal space of the wing transmission system are calculated according to the following formula:

[0020] V i = (V i1 + V i2 ) / m

[0021]

[0022] In the formula, V i1 is the internal space of the wing transmission system, m is a constant, V i2 is the internal space of the balance measurement system, D is the diameter of the circumscribed strut structure, and r is a constant.

[0023] In the deformable aircraft system model for wind tunnel test, the balance measurement system comprises a tightening screw, a heat insulation cone sleeve, a balance and a balance measurement system sealing cover; one end of the balance is sleeved with the heat insulation cone sleeve; the internal part of the aircraft body structure model is connected with the one end of the balance sleeved with the heat insulation cone sleeve through the tightening screw; and the balance measurement system sealing cover is connected with the tail end of the aircraft body structure model through a screw.

[0024] In the deformable aircraft system model for wind tunnel test, the balance measurement system further comprises a compression gasket and a heat insulation gasket; the compression gasket and the heat insulation gasket are both sleeved on the outer surface of the tightening screw; and the compression gasket and the heat insulation gasket are both located between the internal part of the aircraft body structure model and the one end of the balance, the compression gasket is in contact with the internal part of the aircraft body structure model, and the heat insulation gasket is in contact with the one end of the balance.

[0025] In the deformable aircraft system model for wind tunnel test, the wing surface transmission system comprises a motor, a motor heat insulation sleeve, a sliding block, a wing surface transmission system sealing cover, a motor wire slot, a motor fixing cover, a connecting rod and a wing surface; the outer diameter of the motor heat insulation sleeve is matched with the inner wall of the aircraft body structure, and the inner diameter of the motor heat insulation sleeve is matched with the outer diameter of the motor; the motor fixing cover is connected with one end of the motor; the sliding block is sleeved on the outer surface of the output shaft of the motor; one end of the connecting rod is connected with the sliding block through a pin, and the other end of the connecting rod is connected with the wing surface through a pin; and the wing surface transmission system sealing cover is connected with the tail end of the aircraft body structure model through a screw.

[0026] Compared with the prior art, the deformable aircraft system model for wind tunnel test has the following beneficial effects:

[0027] (1) The aircraft internal space is fully utilized, and the wing surface transmission system and the balance measurement system can be simultaneously accommodated in a small aircraft;

[0028] (2) The wing surface transmission system is sealed, so that impurities cannot enter the system and the wing surface movement is not blocked;

[0029] (3) The interference problem of the wing surface transmission system to the balance measurement system is solved, and the accurate measurement of the aerodynamic force of the deformable aircraft is realized;

[0030] (4) The temperature effect of high-temperature gas on the balance measurement system is avoided. BRIEF DESCRIPTION OF DRAWINGS

[0031] Various other advantages and benefits will become apparent to those of ordinary skill in the art upon reading the following detailed description of the preferred embodiments with reference made to the accompanying drawings. The drawings are for purposes of illustration only and are not intended to be limiting in

[0032] Figure 1 is a schematic diagram of a hypersonic wind tunnel test method of a morphing aircraft provided by an embodiment of the present application;

[0033] Figure 2 is a structural schematic diagram of a morphing aircraft system model for a wind tunnel test provided by an embodiment of the present application;

[0034] Figure 3 is another structural schematic diagram of a morphing aircraft system model for a wind tunnel test provided by an embodiment of the present application;

[0035] Figure 4 is a schematic diagram of a wing surface transmission system and a balance measurement system provided by an embodiment of the present application;

[0036] Figure 5 is another schematic diagram of a wing surface transmission system and a balance measurement system provided by an embodiment of the present application;

[0037] Figure 6 is a schematic diagram of a motor fixing cover structure in a wing surface transmission system provided by an embodiment of the present application;

[0038] Figure 7 is a schematic diagram of a balance transmission system sealing cover structure provided by an embodiment of the present application;

[0039] Figure 8 is a schematic diagram of a wing surface transmission system sealing cover structure provided by an embodiment of the present application. DETAILED DESCRIPTION

[0040] Exemplary embodiments of the present disclosure will be described hereinafter with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the present disclosure to those skilled in the art. It should be noted that the embodiments in the present application and the features in the embodiments can be combined with each other without conflict. The present application will be described in detail below with reference to the accompanying drawings and in conjunction with the embodiments.

[0041] Figure 1 is a schematic diagram of a hypersonic wind tunnel test method of a morphing aircraft provided by an embodiment of the present application; Figure 2is a structural schematic diagram of a deformable aircraft system model for a wind tunnel test provided by an embodiment of the present application; Figure 3 is another structural schematic diagram of a deformable aircraft system model for a wind tunnel test provided by an embodiment of the present application. As shown in Figure 1 , Figure 2 and Figure 3 , the deformable aircraft system model for a wind tunnel test comprises an aircraft body structure model 1, an external support pole structure 2, a balance measurement system 3, a wing surface transmission system 4 and an aircraft internal partition 5. Among them,

[0042] One end of the external support pole structure 2 is connected with the tail end of the aircraft body structure model 1, and the other end of the external support pole structure 2 is connected with an angle of attack mechanism; the balance measurement system 3 and the wing surface transmission system 4 are arranged at the rear end inside the aircraft body structure model 1; the balance measurement system 3 is located at the upper part of the wing surface transmission system 4, and the aircraft internal partition 5 is arranged between the balance measurement system 3 and the wing surface transmission system 4.

[0043] Since the actual deformable aircraft is large in size, and the hypersonic wind tunnel is generally small in size due to the structure, the deformable aircraft cannot be directly used for wind tunnel test and must be scaled. The main parameter of the model size limit in the wind tunnel test is the model blockage, and the deformable aircraft model scaling is designed according to the actual situation. After the scaling is determined, the model size is also determined. On this basis, the internal space size of the model is quantitatively analyzed, so the balance measurement system and the wing surface transmission system can be designed and laid out according to the size of the internal cavity. Further, after the internal space system layout of the model is determined, the available space of the two subsystems can be accurately calculated. Finally, the appropriate measurement balance and test motor are selected according to the existing size information and the actual test requirements of the deformable aircraft model.

[0044] Specifically, the scaling of the deformable aircraft system model is obtained according to the model blockage, and the size of the deformable aircraft system model is obtained according to the scaling of the deformable aircraft system model.

[0045] The scaling of the deformable aircraft system model is obtained by the following formula:

[0046] ε=k[1-q(λ) / σ];

[0047] q(λ)=A M / A T ;

[0048] Wherein, A M is the first throat cross-sectional area of the wind tunnel; q(λ) is a gas dynamics function, which can be obtained from the gas flow function table; λ is the wavelength of the gas; σ is the total pressure recovery coefficient of the supersonic normal shock; k is a constant, usually k=0.25-0.40.

[0049] The dimensions of the deformable aircraft system model are:

[0050] V t =V·ε;

[0051] In the formula, V t ε is the size of the deformable aircraft system model, V is the volume of the aircraft, and ε is the scale of the deformable aircraft system model.

[0052] The internal space of the deformable aircraft system model is obtained based on its dimensions.

[0053] The formula for calculating the internal space of a deformable aircraft system model is as follows:

[0054] V i =V t ·n;

[0055] In the formula, V i It is the internal space of the deformable aircraft model; n is a constant, which is related to the shape of the model, and usually n = 0.5 to 0.8.

[0056] The usable space of the balance measurement system 3 and the wing surface transmission system 4 is obtained from the internal space of the deformable aircraft system model.

[0057] The calculation formulas for the space utilization of the balance measurement system 3 and the airfoil transmission system 4 are as follows:

[0058] V i =(V i1 +V i2 ) / m;

[0059]

[0060] In the formula, V i1 It is the internal space of the airfoil drive system; m is a constant, related to the shape of the model, usually m = 0.9 to 0.95; V i2 D is the internal space of the balance measurement system; D is the diameter of the model support rod; r is a constant, which is related to the shape of the model, usually r = 1 to 2.

[0061] Figure 2 and Figure 3The schematic diagram of the deformable aircraft system model for wind tunnel test is shown. The aircraft body structure model 1 is used to realize the aircraft wing deformation and other functions, while providing sufficient internal space to accommodate the balance measurement system and the wing transmission system. The external strut structure 2 realizes the support function of the aircraft. The balance measurement system 3 is used to realize the measurement of the aerodynamic coefficients of the aircraft in the wind tunnel. The wing transmission system 4 is used to realize the deformation of the aircraft wing in the wind tunnel test. The aircraft internal partition 5 is used to separate the wing transmission system and the balance measurement system to avoid mutual interference between the two systems.

[0062] As shown in Figure 4 and Figure 5 , the balance measurement system 3 includes a pull screw 3-1, a compression gasket 3-2, a heat insulation gasket 3-3, a heat insulation cone sleeve 3-4, a balance 3-5 and a balance measurement system sealing cover 3-6; wherein,

[0063] One end of the balance 3-5 is sleeved with the heat insulation cone sleeve 3-4; the internal part of the aircraft body structure model 1 is connected with one end of the balance 3-5 sleeved with the heat insulation cone sleeve 3-4 through the pull screw 3-1; the balance measurement system sealing cover 3-6 is connected with the tail end of the aircraft body structure model 1 through a screw; the compression gasket 3-2 and the heat insulation gasket 3-3 are both sleeved on the outer surface of the pull screw 3-1; the compression gasket 3-2 and the heat insulation gasket 3-3 are both located between the internal part of the aircraft body structure model 1 and one end of the balance 3-5, the compression gasket 3-2 is in contact with the internal part of the aircraft body structure model 1, and the heat insulation gasket 3-3 is in contact with one end of the balance 3-5.

[0064] The pull screw 3-1 cooperates with the balance system to realize the connection of the balance and the deformable aircraft. At the same time, the existence of the compression gasket 3-2 increases the contact area, so that the connection of the two is more close.

[0065] The heat insulation gasket 3-3 and the heat insulation cone sleeve 3-4 are both made of glass steel material, and the main function is to separate the balance and the aircraft structure to avoid direct contact between the two, and to prevent the heat from directly transferring from the aircraft structure to the balance, thereby avoiding the interference of temperature on the system.

[0066] The balance 3-5 is the main component for measuring the aerodynamic coefficients of the deformable aircraft.

[0067] The balance measurement system sealing cover 3-6 has a structure as shown in Figure 7 , and the main function is to prevent the high temperature and high pressure gas in the wind tunnel test from directly contacting the balance. At the same time, the sealing cover is provided with a strut hole and a motor wire hole, and the main function is to facilitate the leading out of the motor wire and the strut.

[0068] As shown in Figure 4 and Figure 5As shown, the aerofoil transmission system 4 includes a motor 4-1, a motor heat insulation sleeve 4-2, a slider 4-3, an aerofoil transmission system sealing cover 4-4, a motor wire slot 4-5, a motor fixing cover 4-6, a connecting rod 4-7 and an aerofoil 4-8; wherein,

[0069] The outer diameter of the motor heat insulation sleeve 4-2 matches the inner wall of the aircraft body structure, and the inner diameter of the motor heat insulation sleeve 4-2 matches the outer diameter of the motor 4-1; the motor fixing cover 4-6 is connected to one end of the motor 4-1; the slider 4-3 is sleeved on the outer surface of the output shaft of the motor 4-1; one end of the connecting rod 4-7 is connected to the slider 4-3 through a pin, and the other end of the connecting rod 4-7 is connected to the aerofoil 4-8 through a pin; the aerofoil transmission system sealing cover 4-4 is connected to the tail end of the aircraft body structure model 1 through a screw

[0070] The motor 4-1, the slider 4-3, the connecting rod 4-7 and the aerofoil 4-8 are the main components of the aerofoil movement, and the main function is to realize the movement of the aerofoil.

[0071] The motor heat insulation sleeve 4-2 is made of glass steel material, and the main function is to separate the motor and the aircraft structure, so as to achieve the purpose of heat insulation and avoid the failure of the motor due to high temperature.

[0072] The aerofoil transmission system sealing cover 4-4, as shown in Figure 8 separates the aerofoil transmission system and the external environment, which can avoid the influence of high temperature and high pressure gas on the motor, and prevent foreign impurities from entering the system, thereby avoiding the system from being stuck.

[0073] The motor fixing cover 4-6, as shown in Figure 6 fixes the motor in the deformable aircraft structure.

[0074] Since the motor is fixed in the aircraft structure in an inverted manner, the leading of the motor wire is very important. The motor wire slot 4-5 is a slot hole parallel to the motor axis, and the main function is to lead the motor wire. The motor wire is led out from the motor wire slot 4-5, then introduced into the balance measurement system cavity, and finally led out to the external environment from the motor wire hole of the balance system sealing cover.

[0075] The present application determines the scaling ratio of the real shape of the morphing aircraft and the wind tunnel model according to the blocking ratio requirement of the hypersonic wind tunnel, thereby determining the size of the morphing aircraft model, and then arranges the wing surface transmission system and the balance measurement system according to the size of the aircraft model, and then selects the measurement balance and the transmission motor with appropriate size according to the requirement. Inside the structure of the morphing aircraft, the balance measurement system and the wing surface transmission system are arranged up and down, and are separated by a partition in the middle, thereby avoiding the measurement interference of the transmission system to the aerodynamic force. In order to reduce the interference of the temperature effect to the balance measurement system and the wing surface transmission system, the following methods are proposed: for the balance measurement system, a heat insulation cone sleeve is designed to avoid the direct transmission of heat energy from the model to the balance; for the wing surface transmission system, a heat insulation motor cylinder is designed to avoid the failure of the motor due to high temperature. At the same time, the balance measurement system and the wing surface transmission system are both designed with a sealing cover to prevent the external airflow and impurities from directly affecting the internal system of the morphing aircraft.

[0076] Although the present application has been disclosed with the above preferred embodiments, it is not intended to limit the present application, and any person skilled in the art can make possible changes and modifications to the technical solutions of the present application by using the disclosed methods and technical contents without departing from the spirit and scope of the present application. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application, which does not deviate from the technical solutions of the present application, all belong to the protection scope of the technical solutions of the present application.

Claims

1. A deformable aircraft system model for wind tunnel testing, characterized in that The application relates to a model of a deformable aircraft system, which comprises an aircraft body structure model (1), an external supporting rod structure (2), a balance measuring system (3), a wing surface transmission system (4) and an aircraft internal partition (5), wherein the tail end of the external supporting rod structure (2) is connected with the tail end of the aircraft body structure model (1), the other end of the external supporting rod structure (2) is connected with an angle of attack mechanism, the balance measuring system (3) and the wing surface transmission system (4) are arranged at the rear end of the inside of the aircraft body structure model (1), the balance measuring system (3) is located at the upper part of the wing surface transmission system (4), and the aircraft internal partition (5) is arranged between the balance measuring system (3) and the wing surface transmission system (4). The tail end of the external supporting rod structure (2) is connected with the tail end of the aircraft body structure model (1), the other end of the external supporting rod structure (2) is connected with an angle of attack mechanism; The balance measuring system (3) and the wing surface transmission system (4) are arranged at the rear end of the inside of the aircraft body structure model (1), the balance measuring system (3) is located at the upper part of the wing surface transmission system (4), and the aircraft internal partition (5) is arranged between the balance measuring system (3) and the wing surface transmission system (4). The scale of the deformable aircraft system model is obtained according to the model block degree, and the scale of the deformable aircraft system model is obtained through the following formula: The scale of the deformable aircraft system model is obtained through the following formula: The internal space of the balance measuring system (3) and the internal space of the wing surface transmission system (4) are obtained according to the internal space of the deformable aircraft system model, and the calculation formula of the internal space of the balance measuring system (3) and the internal space of the wing surface transmission system (4) is as follows: The wing surface transmission system (4) comprises a motor (4-1), a motor heat insulation sleeve (4-2), a sliding block (4-3), a wing surface transmission system sealing cover (4-4), a motor wire slot (4-5), a motor fixed cover (4-6), a connecting rod (4-7) and a wing surface (4-8), wherein the outer diameter of the motor heat insulation sleeve (4-2) is matched with the inner wall of the aircraft body structure, and the inner diameter of the motor heat insulation sleeve (4-2) is matched with the outer diameter of the motor (4-1); The motor fixed cover (4-6) is connected with one end of the motor (4-1); q(λ) = A M / A T ; where A M is the first throat cross-sectional area of the wind tunnel; q(λ) is a gas dynamics function; λ is the gas wavelength; σ is the total pressure recovery coefficient of a normal shock wave at supersonic speed; and k is a constant. The sliding block (4-3) is sleeved on the outer surface of the output shaft of the motor (4-1); V i = (V i1 + V i2 ) / m; where V i1 is the internal space of the wing transmission system; m is a constant; V i2 is the internal space of the balance measurement system; D is the diameter of the external connecting strut structure; r is a constant; One end of the connecting rod (4-7) is connected with the sliding block (4-3) through a pin, and the other end of the connecting rod (4-7) is connected with the wing surface (4-8) through a pin; The wing surface transmission system sealing cover (4-4) is connected with the tail end of the aircraft body structure model (1) through screws. The constant k is 0.25-0.

40. The size of the deformable aircraft system model is as follows: The internal space of the deformable aircraft system model is obtained according to the size of the deformable aircraft system model, and the calculation formula of the internal space of the deformable aircraft system model is as follows: The balance measuring system (3) comprises a tightening screw (3-1), a heat insulation cone sleeve (3-4), a balance (3-5) and a balance measuring system sealing cover (3-6), wherein one end of the balance (3-5) is sleeved with the heat insulation cone sleeve (3-4); 2. The morphing aircraft system model for wind tunnel testing of claim 1, wherein: The internal part of the aircraft body structure model (1) is connected with one end of the balance (3-5) sleeved with the heat insulation cone sleeve (3-4) through the tightening screw (3-1); 3. The morphing aircraft system model for wind tunnel testing of claim 1, wherein: The balance measuring system sealing cover (3-6) is connected with the tail end of the aircraft body structure model (1) through screws. V t = V · ε; where V t is the morphing aircraft system model size, V is the aircraft volume, and ε is the scale of the morphing aircraft system model.

4. The morphing aircraft systems model for wind tunnel testing of claim 1, wherein: ​ V i = V t · n; where V i is the internal space of the morphing aircraft system model; n is a constant.

5. The morphing aircraft systems model for wind tunnel testing of claim 1, wherein: ​ ​ ​ ​ 6. A deformable aircraft systems model for wind tunnel testing according to claim 5, characterized in that: The balance measurement system (3) further comprises a compression gasket (3-2) and a heat insulation gasket (3-3); wherein, The compression gasket (3-2) and the heat insulation gasket (3-3) are both sleeved on the outer surface of the compression screw (3-1); The compression gasket (3-2) and the heat insulation gasket (3-3) are both located between the internal part of the aircraft body structure model (1) and one end of the balance (3-5), the compression gasket (3-2) is in contact with the internal part of the aircraft body structure model (1), and the heat insulation gasket (3-3) is in contact with one end of the balance (3-5).

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