Aerospace vehicle rudder airfoil force and heat combined test method and device
By converting the distributed aerodynamic power on the rudder surface of the aerospace aircraft into concentrated force, loading it at the intersection of the ribs and beams, combining the limit structure and heating device, the problem of load equivalent inaccurate in traditional test methods is solved, and the accuracy and reliability of the test are improved.
Patent Information
- Application Number
- CN202510283046.7
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-03-11
- Publication Date
- 2025-06-03
AI Technical Summary
The combined force-thermal test method of traditional aerospace aircraft has the problem of inaccurate load equivalent, resulting in inaccurate test results and low reliability.
By converting the distributed pneumatic power on the surface of the rudder flange into multiple concentrated forces and loading it at the intersection of the ribs and beams, combining the limit structure to constrain the loading direction, a heating device is used to apply a heating environment to the rudder flange to collect force-heat joint test data.
Ensure the rationality and accuracy of load application, avoid local stress concentration or failure, and improve the accuracy and reliability of the test.
Smart Images
Figure CN120081009A_ABST
Abstract
Description
Technical Field
[0001] This application relates to the technical field of force - heat combined tests for the rudder and wing surfaces of aerospace vehicles, and particularly to a method and device for force - heat combined tests of the rudder and wing surfaces of aerospace vehicles. Background Art
[0002] When an aerospace vehicle travels at high speed through the atmosphere, its external structure is subjected to strong aerodynamic forces and aerodynamic heating. The surface temperature of the rudder and wing surface structure can even reach up to thousands of degrees Celsius. The harsh service environment poses a severe challenge to its structural safety. Ground thermal intensity tests are an important means to determine the safety of the structural design of aerospace vehicles.
[0003] However, in traditional thermal intensity test methods, there are inaccurate temperatures in load equivalence. Existing methods use concentrated forces to replace distributed forces, but the selection principles and loading positions of the concentrated forces are not clearly defined, which easily leads to distortion of the local stress distribution, and further results in inaccurate final test effects and low reliability. Summary of the Invention
[0004] In view of this, this application provides a method and device for force - heat combined tests of the rudder and wing surfaces of aerospace vehicles to solve the problem of low accuracy in existing force - heat combined tests of the rudder and wing surfaces of aerospace vehicles.
[0005] To solve the above - mentioned technical problems, one technical solution adopted in this application is: to provide a method for force - heat combined tests of the rudder and wing surfaces of aerospace vehicles, including: converting the distributed aerodynamic forces on the surface of the rudder and wing surface structure into multiple concentrated forces through the principles of resultant force equivalence and moment equivalence, and the loading areas of the concentrated forces are located at the intersections of the ribs and beams of the rudder and wing surface structure; using a loading device to respectively load the corresponding concentrated forces in each loading area, and the loading device includes a loading rod for transmitting concentrated static loads, a limiting structure for defining the loading direction, and a loading head for fitting with the surface of the rudder and wing surface structure to apply the concentrated force; applying a thermal environment to the surface of the rudder and wing surface structure through a heating device; collecting force - heat combined test data of the rudder and wing surface structure.
[0006] As a further improvement of this application, the heating device includes a radiation heating structure, a thermal compensation structure, a thermocouple, and a controller; applying a thermal environment to the surface of the rudder and wing surface structure through the heating device includes: using the radiation heating structure to heat the non - loading areas on the surface of the rudder and wing surface structure; using the thermocouple to measure the temperature difference between the loading areas and non - loading areas on the surface of the rudder and wing surface structure, where the loading areas are in contact with the loading head and the non - loading areas are not in contact with the loading head; using the controller to confirm whether the temperature difference reaches a preset temperature difference threshold, and when the temperature difference between the loading areas and non - loading areas reaches the preset temperature difference threshold, starting the thermal compensation structure, and when the temperature difference between the loading areas and non - loading areas does not reach the preset temperature difference threshold, closing the thermal compensation structure.
[0007] To solve the above technical problems, another technical solution adopted by this application is: to provide a combined force and heat test device for the rudder wing surface of an aerospace vehicle, which is applied to the combined force and heat test method for the rudder wing surface of the aerospace vehicle as described above; the device includes a loading device and a heating device; the loading device includes a loading rod, a limiting structure, and a loading head. The loading rod is used to transmit a concentrated static load. The limiting structure is sleeved on the loading rod and is used to constrain the loading direction of the loading rod. The loading head is arranged on the loading rod and is used to contact the loading area on the surface of the rudder wing surface structure and apply a concentrated force; the heating device is arranged adjacent to the surface of the rudder wing surface structure of the aerospace vehicle and is used to heat the surface of the rudder wing surface structure of the aerospace vehicle.
[0008] As a further improvement of this application, the heating device includes a radiation heating structure, a thermal compensation structure, a thermocouple, and a controller; the radiation heating structure is arranged adjacent to the non-loading area on the surface of the rudder wing surface structure of the aerospace vehicle and is used to heat the surface of the rudder wing surface structure of the aerospace vehicle; the thermal compensation structure is arranged around the loading head and is used to heat the loading head; the thermocouple is arranged in the loading area and is used to measure the temperature difference between the loading area and the non-loading area; the controller is electrically connected to the thermocouple and the thermal compensation structure respectively, and is used to start the thermal compensation structure when the temperature difference between the loading area and the non-loading area reaches a preset temperature difference threshold, and to turn off the thermal compensation structure when the temperature difference between the loading area and the non-loading area does not reach the preset temperature difference threshold.
[0009] As a further improvement of this application, a high thermal conductivity insulating layer is arranged between the thermocouple and the loading head.
[0010] As a further improvement of this application, it further includes a follow-up structure. The follow-up structure is arranged between the loading rod and the loading head. The follow-up structure is used to generate a morphological change so that the loading head fits the surface of the rudder wing surface structure of the aerospace vehicle when the loading head contacts the loading area.
[0011] As a further improvement of this application, the follow-up structure includes a metal bellows or a thrust ball bearing.
[0012] As a further improvement of this application, the loading head is designed with a profiling structure so that the shape of the loading head coincides with the corresponding loading area on the surface of the rudder wing surface structure of the aerospace vehicle.
[0013] As a further improvement of this application, the radiation heating structure includes a quartz lamp array, which is arranged to cover the non-loading area on the surface of the rudder wing surface structure of the aerospace vehicle.
[0014] As a further improvement of this application, the thermal compensation structure includes an induction coil, which is arranged to surround the loading head.
[0015] The beneficial effects of this application are:
[0016] The method for the combined force and heat test of the rudder wing surface of the aerospace vehicle in this application converts the distributed aerodynamic force on the surface of the rudder wing surface into multiple concentrated forces based on the principles of resultant force equivalence and moment equivalence, and selects the intersection of the rib plate and the beam as the loading position (main load-bearing structure), ensuring the rationality and accuracy of the load application, and avoiding local stress concentration or failure of the structure caused by improper load distribution. Moreover, during the process of applying the concentrated force, the limiting structure of the loading device is used to strictly restrict the loading direction, ensuring that the concentrated force is applied along the set direction, avoiding the instability of the loading direction, and ensuring the accuracy of the final test structure. BRIEF DESCRIPTION OF THE DRAWINGS
[0017] Figure 1 is a schematic structural diagram of an embodiment of the combined force and heat test device for the rudder wing surface of the aerospace vehicle of the present invention;
[0018] Figure 2 is a schematic diagram of the electrical connection relationship of an embodiment of the combined force and heat test device for the rudder wing surface of the aerospace vehicle of the present invention;
[0019] Figure 3 is a schematic flow diagram of an embodiment of the method for the combined force and heat test of the rudder wing surface of the aerospace vehicle of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS
[0020] Next, the technical solutions in the embodiments of the present application will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present application.
[0021] The terms "first", "second", and "third" in this application are only used for descriptive purposes, and cannot be understood as indicating or implying relative importance or implicitly specifying the quantity of the indicated technical features. Thus, the features defined with "first", "second", and "third" may explicitly or implicitly include at least one of such features. In the description of this application, the meaning of "a plurality" is at least two, such as two, three, etc., unless otherwise specifically and clearly defined. All directional indications (such as up, down, left, right, front, back...) in the embodiments of this application are only used to explain the relative spatial position and movement conditions between components in a specific posture (as shown in the drawings). If the specific posture changes, the directional indications will also change accordingly. In addition, the terms "include" and "have" and any variations thereof are intended to cover non-exclusive inclusion. For example, a process, method, device, product, or equipment that includes a series of steps or units is not limited to the listed steps or units, but optionally further includes steps or units not listed, or optionally further includes other steps or units inherent to these processes, methods, products, or equipment.
[0022] As used herein, the mention of "embodiments" means that the specific features, structures, or characteristics described in connection with the embodiments may be included in at least one embodiment of the present application. The phrase appears in various places in the specification and does not necessarily refer to the same embodiment, nor is it an independent or alternative embodiment mutually exclusive with other embodiments. Those skilled in the art will explicitly and implicitly understand that the embodiments described herein may be combined with other embodiments.
[0023] Figure 1 is a schematic structural diagram of the force-thermal combined test device for the rudder wing surface of an aerospace vehicle according to an embodiment of the present invention. As Figure 1 shown, the force-thermal combined test device for the rudder wing surface of the aerospace vehicle includes a loading device 1 and a heating device 2.
[0024] The loading device 1 includes a loading rod 11, a limiting structure 12, and a loading head 13. Among them, the loading rod 11 is used to transmit a concentrated static load. The limiting structure 12 is sleeved on the loading rod 11 and is used to constrain the loading direction of the loading rod 11. The loading head 13 is arranged on the loading rod 11 and is used to contact the loading area on the surface of the rudder wing surface structure and apply a concentrated force. It should be noted that the loading head 13 is made of a high-temperature-resistant material, and generally a metal material can be selected.
[0025] The heating device 2 is arranged adjacent to the surface of the rudder wing surface structure of the aerospace vehicle and is used to heat the surface of the rudder wing surface structure of the aerospace vehicle. The heating method can be achieved by radiation heating.
[0026] Specifically, when using the force-thermal combined test device for the rudder wing surface of the aerospace vehicle to conduct a force-thermal combined test on the rudder wing surface of the aerospace vehicle, after confirming the loading area on the surface of the rudder wing surface structure, a loading device 1 is correspondingly arranged in each loading area, and the heating device 2 is arranged on the surface of the rudder wing surface structure. Then, after aligning the loading head 13 with the loading area, the loading head 13 is brought into contact with the loading area, and at the same time, the loading direction of the loading rod 11 is limited by the limiting structure 12 to prevent the loading rod 11 from tilting or moving, which may affect the loading of the concentrated force. Then, the heating device 2 is turned on to heat the surface of the rudder wing surface structure, and then the force-thermal combined test data of the rudder wing surface structure are collected.
[0027] Further, please refer to Figure 2 together. The heating device 2 includes a radiation heating structure 21, a thermal compensation structure 22, a thermocouple 23 ( Figure 1 not shown in Figure 1 ) and a controller 24 (
[0028] Among them, the radiation heating structure 21 is arranged adjacent to the unloaded area on the surface of the aeronautical and aerospace vehicle rudder wing surface structure for heating the surface of the aeronautical and aerospace vehicle rudder wing surface structure. The thermal compensation structure 22 is arranged around the loading head 13 for heating the loading head 13. The thermocouple 23 is arranged in the loading area for measuring the temperature difference between the loading area and the unloaded area. The controller 24 is electrically connected to the thermocouple 23 and the thermal compensation structure 22 respectively, and is used to start the thermal compensation structure 22 when the temperature difference between the loading area and the unloaded area reaches the preset temperature difference threshold, and turn off the thermal compensation structure 22 when the temperature difference between the loading area and the unloaded area does not reach the preset temperature difference threshold.
[0029] It should be noted that during the test process, it usually occurs that the temperature of the area contacted by the loading head 13 is significantly lower than the surface of the structure that is not covered by the loading head 13. In the traditional test method, the low-temperature area is not thermally compensated, resulting in inaccurate measurement results.
[0030] In this embodiment, by arranging the thermal compensation structure 22 around the loading head 13, the thermal compensation structure 22 can heat the loading head 13, and then transfer the heat to the loading area through the loading head 13, so as to thermally compensate the loading area, avoid the situation of too large temperature difference between the loading area and the unloaded area, and further improve the accuracy and reliability of the force-thermal combined test results. Specifically, a preset temperature difference threshold (calculated according to the structural thermal stress standard) is preset in this embodiment. When the temperature difference measured by the thermocouple 23 between the loading area and the unloaded area exceeds the preset temperature difference threshold, for example, exceeds 100 °C or the relative temperature difference exceeds 30%, the controller 24 controls the thermal compensation device to turn on and heat the loading head 13, and the loading head 13 transfers the heat to the loading area, so as to thermally compensate the loading area; when the temperature difference measured by the thermocouple 23 between the loading area and the unloaded area does not reach the preset temperature difference threshold, thermal compensation is not required and the thermal compensation device is turned off.
[0031] Furthermore, a high thermal conductivity insulating layer (not shown in the figure) is arranged between the thermocouple 23 and the loading head 13 to insulate the thermocouple 23 from the loading head 13 and ensure that the measured temperature is the surface temperature of the rudder wing surface structure.
[0032] Furthermore, the force-thermal combined test device for the aeronautical and aerospace vehicle rudder wing surface further includes a follow-up structure 14. The follow-up structure 14 is arranged between the loading rod 11 and the loading head 13, and is used to generate a morphological change so that the loading head 13 fits with the surface of the aeronautical and aerospace vehicle rudder wing surface structure when the loading head 13 contacts in the loading area.
[0033] It should be noted that the surface of the rudder wing surface structure is irregular, and some areas may not be able to fit well with the loading head 13, resulting in poor transfer effect of the concentrated force and affecting the accuracy of the test results.
[0034] In this embodiment, in order to better make the loading head 13 fit the surface of the rudder wing surface structure, a follow-up structure 14 is provided. The follow-up structure 14 does not limit the direction of the loading head 13, so that the loading head 13 can change its direction following the shape of the surface of the rudder wing surface structure, making the loading head 13 fit the surface of the rudder wing surface structure more closely.
[0035] Furthermore, the follow-up structure 14 includes a metal bellows or a thrust ball bearing. Among them, the specifications of the metal bellows are selected according to the magnitude of the concentrated force at this point.
[0036] Furthermore, the loading head 13 is designed with a profiling structure, so that the shape of the loading area corresponding to the surface of the rudder wing surface structure of the aerospace vehicle coincides with each other.
[0037] Specifically, the loading head 13 in contact with the rudder wing surface structure of the loading device 1 is designed with a profiling structure to achieve the maximum contact area.
[0038] Furthermore, the radiation heating structure 21 includes a quartz lamp array, which is arranged to cover the non-loading area on the surface of the rudder wing surface structure of the aerospace vehicle.
[0039] Specifically, according to the shape of the rudder wing surface structure, the radiation heating structure 21 is arranged in the area outside the interference influence area of the loading device 1. It can be understood that the radiation heating structure 21 should be electrically insulated from other devices.
[0040] Furthermore, the thermal compensation structure 22 includes an induction coil (not shown in the figure), and the induction coil is arranged around the loading head 13.
[0041] Specifically, an induction coil is designed around the circumference of the metal loading head 13. The induction coil is connected to the controller 24. During the test, local heating of the loading head 13 is achieved through electromagnetic induction to compensate for the blocked radiation heating part and reduce the temperature difference between the loading area and the non-loading area of the rudder wing surface structure.
[0042] Figure 3 Shows a schematic flow chart of the force-thermal combined test method for the rudder wing surface of the aerospace vehicle in the embodiment of the present invention. As Figure 3 shown, the force-thermal combined test method for the rudder wing surface of the aerospace vehicle includes:
[0043] Step S1: Convert the distributed aerodynamic force on the surface of the rudder wing surface structure into multiple concentrated forces through the principles of resultant force equivalence and moment equivalence. The loading area of the concentrated force is located at the intersection of the rib and the beam of the rudder wing surface structure.
[0044] Specifically, in the thermal strength test of the rudder wing surface of the aerospace vehicle, the distributed aerodynamic force (such as aerodynamic pressure) needs to be equivalently converted into several concentrated forces to ensure the feasibility and accuracy of the loading. The core principle of the equivalent conversion is:
[0045] Equivalent resultant force: The sum of the transformed concentrated forces is equal to the total resultant force of the original distributed force;
[0046] Equivalent moment: The total moment of the transformed concentrated forces about the reference point is equal to the total moment of the original distributed force about the same reference point.
[0047] Based on this, a mathematical model is established:
[0048] 1. Define the distributed aerodynamic force function
[0049] Let the distributed aerodynamic force on the surface of the rudder wing be q(x, y) (force per unit area, unit: N / m 2 ), and the projection area of the rudder wing be S. The total resultant force F total and the total moment M total are respectively:
[0050] F total = ∫∫q(x, y)dS;
[0051]
[0052] where is the position vector from the action point of the distributed force to the reference point.
[0053] 2. Select the concentrated force loading points
[0054] According to the structural characteristics, select the intersection points of the internal ribs and the beam of the rudder wing as the concentrated force loading positions (denoted as P 1 , P 2 , …, P n ), and these points are usually the main load-bearing nodes and can effectively transfer the load.
[0055] 3. Establish the equivalent equations
[0056] Assume that the magnitudes of the transformed concentrated forces are F 1 , F 2 , …, F n , and the action points are P 1 , P 2 , …, P n , then it is necessary to satisfy:
[0057] (Equivalent resultant force);
[0058] (Equivalent moment);
[0059] where is the position vector from the i-th concentrated force action point to the reference point.
[0060] 4. Solve for the magnitudes and directions of the concentrated forces
[0061] Simplification of symmetric structure: If the rudder wing surface is of symmetric structure, it can be assumed that the magnitudes of the concentrated forces at symmetric positions are equal, reducing the number of unknowns.
[0062] Numerical solution: Determine the magnitudes of each concentrated force through linear algebra methods (such as matrix solution). If the number of equations is insufficient (e.g., the number of unknowns is more than the number of equations), it is necessary to distribute the load in combination with the structural bearing capacity (such as distributing according to the stiffness ratio of rib plates).
[0063] Example: For a rectangular rudder wing surface, if 4 symmetric loading points P 1 ~, P 4 are selected, it can be set that F 1 = F 2 = F 3 = F 4 = F. Then the total resultant force is 4F = F total , and by solving, F = F total / 4. It is necessary to verify whether the moment equivalence is satisfied.
[0064] Step S2: Use a loading device to apply the corresponding concentrated forces in each loading area respectively. The loading device includes a loading rod for transmitting the concentrated static load, a limiting structure for defining the loading direction, and a loading head for fitting with the surface of the rudder wing surface structure to apply the concentrated force.
[0065] Step S3: Apply a thermal environment to the surface of the rudder wing surface structure through a heating device.
[0066] Step S4: Collect the force-thermal coupling test data of the rudder wing surface structure.
[0067] In this embodiment, by transforming the distributed aerodynamic force on the surface of the rudder wing surface into multiple concentrated forces based on the principles of resultant force equivalence and moment equivalence, and selecting the intersection of the rib plate and the beam as the loading position (main load-bearing structure), the rationality and accuracy of the load application are ensured, avoiding local stress concentration or failure of the structure caused by improper load distribution. Moreover, during the process of applying the concentrated force, the limiting structure of the loading device is used to strictly restrict the loading direction, ensuring that the concentrated force is applied along the set direction, avoiding instability of the loading direction, and ensuring the accuracy of the final test structure.
[0068] Furthermore, the heating device includes a radiation heating structure, a thermal compensation structure, a thermocouple, and a controller; Step S3 specifically includes:
[0069] 1. Use the radiation heating structure to heat the non-loading area on the surface of the rudder wing surface structure;
[0070] 2. Use the thermocouple to measure the temperature difference between the loading area and the non-loading area on the surface of the rudder wing surface structure. The loading area is in contact with the loading head, and the non-loading area is not in contact with the loading head;
[0071] 3. Use the controller to confirm whether the temperature difference reaches the preset temperature difference threshold. When the temperature difference between the loading area and the non-loading area reaches the preset temperature difference threshold, start the thermal compensation structure. When the temperature difference between the loading area and the non-loading area does not reach the preset temperature difference threshold, turn off the thermal compensation structure.
[0072] In this embodiment, the temperature difference between the loading area and the non-loading area is measured by a thermocouple. When the temperature difference exceeds the preset temperature difference threshold, the loading head is heated by the thermal compensation structure, and the loading area is thermally compensated by the loading head, thereby reducing the temperature difference between the loading area and the non-loading area. Moreover, when the temperature difference between the loading area and the non-loading area does not reach the preset threshold, there is no need to turn on the thermal compensation structure, reducing energy consumption.
[0073] The above are only the implementation manners of the present application, and do not limit the patent scope of the present application. Any equivalent structure or equivalent process transformation made by using the content of the specification and drawings of the present application, or directly or indirectly applied to other related technical fields, shall be included in the patent protection scope of the present application by the same token.
Claims
1. A combined mechanical and thermal test method for aerospace vehicle rudder wing surfaces, characterized in that: include: The distributed aerodynamic force on the surface of the rudder wing structure is converted into multiple concentrated forces through the principle of equivalent resultant force and equivalent moment, and the loading area of the concentrated force is located at the intersection of the rib plate and the beam of the rudder wing structure; A loading device is used to load the corresponding concentrated force in each loading area, wherein the loading device includes a loading rod for transmitting the concentrated static load, a limiting structure for limiting the loading direction, and a loading head for fitting with the surface of the rudder wing structure to apply the concentrated force; Applying a thermal environment to the surface of the rudder wing structure by means of a heating device; Collect the combined mechanical and thermal test data of the rudder and wing surface structure.
2. The combined mechanical and thermal test method for aerospace vehicle rudder wing surfaces according to claim 1, characterized in that: The heating device comprises a radiation heating structure, a thermal compensation structure, a thermocouple and a controller; applying a thermal environment to the surface of the rudder wing structure through the heating device comprises: Using the radiation heating structure to heat the non-loaded area on the surface of the rudder wing structure; Measuring the temperature difference between a loading area and a non-loading area on the surface of the rudder wing structure by using a thermocouple, wherein the loading area is in contact with the loading head and the non-loading area is not in contact with the loading head; A controller is used to confirm whether the temperature difference reaches a preset temperature difference threshold, and when the temperature difference between the loading area and the non-loading area reaches the preset temperature difference threshold, the thermal compensation structure is started, and when the temperature difference between the loading area and the non-loading area does not reach the preset temperature difference threshold, the thermal compensation structure is turned off.
3. A combined mechanical and thermal test device for rudder and wing surfaces of aerospace vehicles, characterized in that: It is applied to the combined mechanical and thermal test method for the rudder and wing surface of an aerospace vehicle as described in any one of claims 1-2; the device comprises a loading device and a heating device; The loading device comprises a loading rod, a limiting structure and a loading head, wherein the loading rod is used to transmit a concentrated static load, the limiting structure is sleeved on the loading rod and is used to constrain the loading direction of the loading rod, and the loading head is arranged on the loading rod and is used to contact the loading area on the surface of the rudder wing structure and apply a concentrated force; The heating device is disposed adjacent to the surface of the rudder and wing structure of the aerospace vehicle, and is used to heat the surface of the rudder and wing structure of the aerospace vehicle.
4. The aerospace vehicle rudder and wing surface combined mechanical and thermal testing device according to claim 3 is characterized in that: The heating device comprises a radiation heating structure, a thermal compensation structure, a thermocouple and a controller; The radiation heating structure is disposed adjacent to a non-loaded area on the surface of the rudder and wing structure of the aerospace vehicle, and is used to heat the surface of the rudder and wing structure of the aerospace vehicle; The thermal compensation structure is arranged around the loading head and is used for heating the loading head; The thermocouple is arranged in the loading area and is used to measure the temperature difference between the loading area and the non-loading area; The controller is electrically connected to the thermocouple and the thermal compensation structure, respectively, and is used to start the thermal compensation structure when the temperature difference between the loading area and the non-loading area reaches a preset temperature difference threshold, and to shut down the thermal compensation structure when the temperature difference between the loading area and the non-loading area does not reach the preset temperature difference threshold.
5. The aerospace vehicle rudder and wing surface combined mechanical and thermal testing device according to claim 4, characterized in that: A high thermal conductivity insulation layer is arranged between the thermocouple and the loading head.
6. The aerospace vehicle rudder and wing surface combined mechanical and thermal testing device according to claim 3, characterized in that: It also includes a follower structure, which is arranged between the loading rod and the loading head. The follower structure is used to produce a morphological change when the loading head contacts the loading area so that the loading head fits the surface of the rudder wing structure of the aerospace vehicle.
7. The aerospace vehicle rudder and wing surface combined mechanical and thermal testing device according to claim 6, characterized in that: The follower structure includes a metal bellows or a thrust ball bearing.
8. The aerospace vehicle rudder and wing surface combined mechanical and thermal testing device according to claim 3, characterized in that: The loading head adopts a contoured structure design so that the loading head and the loading area corresponding to the surface of the rudder wing structure of the aerospace vehicle are consistent with each other.
9. The aerospace vehicle rudder and wing surface combined mechanical and thermal testing device according to claim 4, characterized in that: The radiation heating structure includes a quartz lamp array, which is arranged to cover a non-loaded area on the surface of the rudder wing structure of the aerospace vehicle.
10. The aerospace vehicle rudder and wing surface combined mechanical and thermal testing device according to claim 4, characterized in that: The thermal compensation structure includes an induction coil, and the induction coil is arranged around the loading head.