Thermal mechanical fatigue testing device and testing method for rocket thrust chamber material structure

Abstract

The present application relates to the field of aerospace technology, in particular to a thermal engine fatigue testing device and a testing method for a material structure of a rocket thrust chamber; the thermal engine fatigue testing device for a material structure of a rocket thrust chamber comprises a sealed shell and a clamping member, a temperature measuring assembly, and a heating assembly disposed in the sealed shell; the clamping member has at least two clamping parts, with a clamping space formed between adjacent clamping parts, the clamping space being used to clamp a test piece; the clamping part and the test piece have the same curved surface, so that the adjacent clamping parts and the test piece clamped between the adjacent clamping parts can simulate at least part of the convergence section of the thrust chamber. In summary, the test piece needs to be replaced each time a test is conducted due to different thrust chamber structures; in actual use, the sealed shell is in an environment filled with protective gas, the test piece is heated by a heating assembly, and a coolant is introduced into the first cooling channel to simulate a cooling effect, and cyclic loading simulates the inner wall destruction process of a real thrust chamber.

Description

Technical Field

[0001] The present application relates to the field of aerospace technology, and in particular to a thermomechanical fatigue testing device and testing method for rocket thrust chamber material structures. Background Art

[0002] Rockets are one of the main ways for humans to enter space, and their research and development progress has become one of the key contents of various countries. As the "heart" of the rocket - the engine, its development is of paramount importance.

[0003] Taking liquid rockets as an example, due to the extreme environment of high temperature and high pressure in the engine thrust chamber, it is difficult for ordinary metals to perform effectively. Even tungsten with a melting point of 3380℃ may be melted. Therefore, the thermal protection method of "regenerative cooling" is often used to reduce the temperature of the metal material on the thrust chamber wall. Figure 1 As shown, before entering the thrust chamber 1' (the thrust chamber 1' has a convergent section 5') for combustion, the low-temperature propellant stored in the liquid collector 4' flows as a coolant through the cooling channel 3' formed by the inner wall 6' and the outer wall 7' in the thrust chamber, performing convection cooling on the inner wall 6' of the thrust chamber, and at the same time preheating and "regenerating" more energy.

[0004] The inner wall of a regeneratively cooled thrust chamber is typically made of a highly thermally conductive copper-based alloy, only 1-2 mm thick, to facilitate rapid temperature transfer between the combustion and coolant sides. Furthermore, the internal and external temperatures are extreme, with a large temperature differential: the combustion chamber temperature ranges from 2000-3000 Kelvin, while the coolant temperature is only 22-300 Kelvin. Furthermore, the internal and external pressures are high, with a significant pressure differential. The pressure inside the combustion chamber can reach 20 MPa, while the pressure on the coolant side can even reach 30 MPa. Under these conditions, the material structure of the inner wall is susceptible to low-cycle fatigue, ratcheting, high-temperature creep, and ductile fracture. The inner wall gradually bulges and thins along the centerline of the cooling channel until a crack forms, penetrating the combustion chamber and the coolant channel. This failure mechanism is known as the "doghouse effect."

[0005] To achieve safe and reusable rockets, it is necessary to understand the damage mechanism of the thrust chamber material, determine its failure patterns, and determine its service life and reliability. This, in turn, improves the design of the regenerative cooling channels. All of this requires data from ground tests to guide design and simulation.

[0006] Currently, there are three common testing methods:

[0007] (1) The ground test method is a material testing method. A testing machine is used to test the mechanical properties (strength, stiffness, hardness, etc.), physical properties (specific heat capacity, thermal conductivity, electrical conductivity, etc.), and chemical properties of standard test pieces. This method can obtain the performance parameters of the material under specific environmental or operating conditions. Although the testing cost of this method is relatively low, it can only reflect the performance of the material and cannot verify the response of the specific thrust chamber structure composed of the material. In addition, material testing is difficult to achieve the extreme environmental conditions in the actual thrust chamber, that is, it cannot fully simulate the real situation.

[0008] (2) Thrust chamber hot test method. This method is to introduce real process materials into the assembled thrust chamber according to the design documents and perform the expected experimental operations to obtain the actual working parameters of the component when tested under predetermined conditions. For this method: 1) There are few temperature measurement points on the inner wall of the thrust chamber, which cannot form a continuous field, and the measurement points are a certain distance away from the inner wall. It is an indirect measurement with a certain error. 2) The values ​​of its heat flux density and other values ​​come from reverse calculations, that is, they cannot be directly measured. 3) Although the interior of the thrust chamber can meet the requirements of real gas heating and simulate real environmental conditions, it is also difficult to make the external chamber pressure of the thrust chamber reach the real high pressure conditions during the rocket launch process. 4) During the hot test, the deformation of the inner wall cannot be measured, and the initial damage and evolution process cannot be seen. Only the damaged test piece can be analyzed. 5) The hot test is a large-scale test and its cost is very high.

[0009] (3) Thermal engine fatigue panel test method: The test panel is designed based on the dimensional parameters of the thrust chamber throat. Laser heating is used to simulate actual gas heating. The test device is used to explore the doghouse destruction mechanism, facilitating the design and test verification process. For this method: 1) For thermal engine fatigue panel testing, although the doghouse destruction process can be achieved, due to the low absorption rate of the thrust chamber material to laser, a special absorption coating needs to be applied to the heating area, which will affect the deformation of the structure during heating. Due to the limitations of laser heating, the test panel does not consider the influence of the following actual factors: 2) The axial Laval curve curvature of the thrust chamber and the circumferential circular curvature. 3) The mutual extrusion of the circumferential structure.

[0010] Therefore, there is an urgent need for a thermal mechanical fatigue testing device and testing method for rocket thrust chamber material structures to solve the technical problems existing in the prior art to a certain extent. Summary of the Invention

[0011] The purpose of this application is to provide a thermal mechanical fatigue testing device and testing method for rocket thrust chamber material structures, so as to solve, to a certain extent, the technical problem in the prior art that it is impossible to verify the response of a specific thrust chamber structure composed of the material.

[0012] The present application provides a thermomechanical fatigue testing device and testing method for a rocket thrust chamber material structure, comprising a sealed shell and a clamping member, a temperature measuring assembly, and a heating assembly disposed within the sealed shell;

[0013] The clamping member has at least two clamping parts, with a clamping space formed between adjacent clamping parts, the clamping space being used to clamp the test piece; the clamping parts and the test piece have the same curved surface, so that the adjacent clamping parts and the test piece clamped between the adjacent clamping parts can simulate at least a portion of the convergent section of the thrust chamber;

[0014] The heating component faces the test piece and is used to heat the test piece; the temperature measuring component faces the test piece and is used to detect the temperature of the test piece; a first cooling channel for passing a coolant is opened along the side wall of the test piece along its extension direction.

[0015] In the above technical solution, further, the clamping member includes a clamping piece and an end seat;

[0016] There are multiple clamping members, and the clamping space is formed between adjacent clamping members;

[0017] The end seat is arranged at an end of the clamping member away from the object to be tested.

[0018] In the above technical solution, further, a second cooling channel is opened on the side wall of the clamping member along its extension direction; the second cooling channel is used to simulate the cooling channel of the thrust chamber.

[0019] In the above technical solution, further, the clamping member also includes a support base;

[0020] The support base is disposed between adjacent clamping members and is used to support the test piece, so that the test piece and its adjacent clamping members can simulate at least a portion of the convergent section of the thrust chamber.

[0021] In the above technical solution, further, the temperature measurement component is an infrared thermometer.

[0022] In the above technical solution, further, the heating component includes an electromagnetic coil, and the electromagnetic coil can perform single-sided inductive heating on the surface of the test piece.

[0023] The present application also provides a method for thermomechanical fatigue testing of a rocket thrust chamber material structure, comprising the following steps:

[0024] Preparation steps: fix the test piece with a clamping member; arrange the infrared thermometer and the electromagnetic coil above the test piece with intervals and both facing the test piece; fill the sealed shell with protective gas;

[0025] Pre-cooling step: turning on the infrared thermometer and introducing coolant into the first cooling channel and the second cooling channel;

[0026] Thermal test step: After precooling for a first preset time, the heating component is turned on to heat the test piece and allow the test piece to reach the target temperature from the original temperature;

[0027] Post-cooling step: After maintaining the target temperature for a second preset time, the heating component is turned off, and the temperature of the test piece is reduced to the original temperature after a third preset time;

[0028] Relaxation step: stop supplying the coolant to the first cooling channel, and wait for a fourth preset time until the surface temperature of the test piece tends to be stable, thereby completing one test cycle of the test piece.

[0029] In the above technical solution, further, the pre-cooling step further includes the following steps:

[0030] Introducing step: introducing coolant into the first cooling channel through the input pipe, and using a mass flow meter, a thermocouple, and a pressure gauge to respectively measure the mass flow, temperature, and pressure of the coolant input into the test piece;

[0031] Cooling step: a coolant flows through a first cooling channel of the test piece to pre-cool the test piece;

[0032] Output step: The coolant flowing through the first pipe flows out through the output pipe, and the mass flow rate, temperature and pressure of the coolant output to the test piece are measured respectively by using a mass flow meter, a thermocouple and a pressure gauge.

[0033] In the above technical solution, further, the relaxation step further includes the following steps:

[0034] Cycle steps: Repeat the preparation step, pre-cooling step, hot test step, post-cooling step and relaxation step N times in sequence, and observe the surface changes of the test piece.

[0035] In the above technical solution, further, the coolant is liquid nitrogen.

[0036] Compared with the prior art, the present invention has the following advantages:

[0037] The present application provides a thermomechanical fatigue testing device for a rocket thrust chamber material structure, comprising a sealed shell and a clamping member, a temperature measuring assembly, and a heating assembly disposed within the sealed shell;

[0038] The clamping member has at least two clamping parts, with a clamping space formed between adjacent clamping parts, the clamping space being used to clamp the test piece; the clamping parts and the test piece have the same curved surface, so that the adjacent clamping parts and the test piece clamped between the adjacent clamping parts can simulate at least a portion of the convergent section of the thrust chamber;

[0039] The heating component faces the test piece and is used to heat the test piece; the temperature measuring component faces the test piece and is used to detect the temperature of the test piece; a first cooling channel for passing a coolant is opened along the side wall of the test piece along its extension direction.

[0040] Specifically, in summary, the test piece needs to be replaced in each test due to different thrust chamber structures; in actual use, the sealed shell is filled with protective gas, the test piece is heated by the heating component, and coolant is introduced into the first cooling channel to simulate the cooling effect, and cyclic loading simulates the inner wall destruction process of the real thrust chamber.

[0041] The present application also provides a method for thermomechanical fatigue testing of a rocket thrust chamber material structure, comprising the following steps:

[0042] Preparation steps: fix the test piece with a clamping member; arrange the infrared thermometer and the electromagnetic coil above the test piece with intervals and both facing the test piece; fill the sealed shell with protective gas;

[0043] Pre-cooling step: turn on the infrared thermometer and introduce coolant into the first cooling channel;

[0044] Thermal test step: After precooling for a first preset time, the heating component is turned on to heat the test piece and allow the test piece to reach the target temperature from the original temperature;

[0045] Post-cooling step: After maintaining the target temperature for a second preset time, the heating component is turned off, and the temperature of the test piece is reduced to the original temperature after a third preset time;

[0046] Relaxation step: stop supplying the coolant to the first cooling channel, and wait for a fourth preset time until the surface temperature of the test piece tends to be stable, thereby completing one test cycle of the test piece.

[0047] In summary, the present application simulates a part of the thrust chamber convergence section to replace the entire thrust chamber, uses a heating component to heat the test piece to meet the high temperature conditions under the action of real gas, and uses safe and controllable liquid nitrogen to cool the test piece to simulate the thermal protection measures of the real situation. The heating process of the single-sided electromagnetic coil can reflect the changes in the surface temperature of the test piece. BRIEF DESCRIPTION OF THE DRAWINGS

[0048] In order to more clearly illustrate the specific implementation methods of the present application or the technical solutions in the prior art, the following is a brief introduction to the drawings required for use in the specific implementation methods or the description of the prior art. Obviously, the drawings described below are some implementation methods of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without any creative work.

[0049] Figure 1 It is a structural diagram in the prior art;

[0050] Figure 2 A schematic structural diagram of a thermomechanical fatigue testing device for a rocket thrust chamber material structure provided in Example 1 of the present application, viewed from a first perspective;

[0051] Figure 3 A schematic structural diagram of the thermomechanical fatigue testing device for a rocket thrust chamber material structure provided in Example 1 of the present application, viewed from a second perspective;

[0052] Figure 4 This is a schematic structural diagram of the support base in the thermal mechanical fatigue testing device for the rocket thrust chamber material structure provided in Example 1 of the present application;

[0053] Figure 5 A schematic diagram of the parameter structure of the first cooling channel in the thermomechanical fatigue testing method for a rocket thrust chamber material structure provided in Example 2 of the present application;

[0054] Figure 6 A structural schematic diagram of the thermo-mechanical fatigue testing method for rocket thrust chamber material structure provided in Example 2 of the present application.

[0055] Reference numerals:

[0056] 1'-thrust chamber; 3'-cooling channel; 4'-liquid collector; 5'-convergent section; 6'-inner wall; 7'-outer wall; 1-sealed shell; 2-test piece; 3-thrust chamber; 4-first cooling channel; 5-clamping member; 6-end seat; 7-second cooling channel; 8-support base; 9-infrared thermometer; 10-electromagnetic coil; 11-input pipe; 12-output pipe; 13-inner wall; 14-outer wall; 15-gas side; 16-atmospheric pressure side; 17-clamping member; 18-rib. DETAILED DESCRIPTION

[0057] The technical solution of the present application will be clearly and completely described below in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments.

[0058] The components of the embodiments of the present application generally described and shown in the drawings herein may be arranged and designed in a variety of different configurations. Therefore, the following detailed description of the embodiments of the present application provided in the drawings is not intended to limit the scope of the claimed application, but merely represents selected embodiments of the present application.

[0059] Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making any creative work shall fall within the scope of protection of this application.

[0060] In the description of this application, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings and are intended solely to facilitate the description of this application and simplify the description. They do not indicate or imply that the devices or components referred to must have a specific orientation, be constructed, or operate in a specific orientation. Therefore, they should not be construed as limitations on this application. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0061] In the description of this application, it should be noted that, unless otherwise expressly specified or limited, the terms "mounted," "connected," and "connected" should be understood in a broad sense. For example, they can refer to fixed connections, detachable connections, or integral connections; mechanical connections or electrical connections; direct connections or indirect connections through an intermediate medium; and internal connections between two components. Those skilled in the art will understand the specific meanings of the above terms in this application based on the specific circumstances.

[0062] Example 1

[0063] Refer to the following Figures 2 to 4 The present invention describes a thermomechanical fatigue testing device for a rocket thrust chamber material structure provided in accordance with the present application.

[0064] Specifically, the thermomechanical fatigue testing device for the material structure of the rocket thrust chamber includes a sealed shell 1 and a clamping member 17, a temperature measuring component and a heating component arranged in the sealed shell 1;

[0065] More specifically, the clamping member 17 has two clamping portions, with a clamping space formed between adjacent clamping portions for clamping the test piece 2. The clamping portions and the test piece 2 have the same curved surface, allowing the adjacent clamping portions and the test piece 2 clamped therebetween to simulate at least a portion of the convergent section of the thrust chamber 3. A first cooling channel 4 for coolant is defined along the sidewall of the test piece 2 along its extension. Furthermore, the clamping portions and the test piece have the same cooling channel configuration.

[0066] More specifically, the heating assembly faces the test piece 2 and is used to heat the test piece 2 . Preferably, the heating assembly includes an electromagnetic coil 10 , which can perform single-sided inductive heating on the surface of the test piece 2 .

[0067] More specifically, the temperature measuring component faces the test piece 2 and is used to detect the temperature of the test piece 2 ; preferably, the temperature measuring component is an infrared thermometer 9 .

[0068] In summary, the test piece 2 needs to be replaced during each test due to the different thrust chamber 3 structures. In actual use, the sealed shell 1 is heated by the heating component in an environment filled with protective gas, and coolant is introduced into the first cooling channel to simulate the cooling effect. The cyclic loading simulates the inner wall destruction process of the real thrust chamber.

[0069] Specifically, the clamping member 17 includes a clamping member 5 and an end seat 6. Two clamping members 5 are provided, with a clamping space formed between adjacent clamping members 5. The end seat 6 is provided at the end of the clamping member 5 away from the test object 2. More specifically, the end seat 6 facilitates adjustment of the clamping members 5, allowing the two clamping members 5 to clamp the test object 2.

[0070] More specifically, the clamping member 17 further includes a driving member, such as a driving motor, the output shaft of the driving motor is connected to the end seat 6 , and the driving motor drives the end seat 6 so that the clamping member 5 can clamp the test piece 2 .

[0071] More specifically, the clamping member 5 simulates a convergent section of a portion of the thrust chamber 3 , that is, its wall surface is an arc-shaped structure along the circumferential direction and a tapered structure along the axial direction.

[0072] Specifically, a second cooling channel 7 is defined along the sidewall of the clamping member 5 along its extension direction; this second cooling channel 7 is used to simulate the cooling channel of the thrust chamber 3. In actual use, no coolant flows through the second cooling channel 7. The purpose of providing the second cooling channel 7 is to simulate the structure of a real thrust chamber 3, thereby ensuring a more realistic understanding of the structural response of the test piece 2.

[0073] Specifically, the clamping member 17 also includes a support base 8; more specifically, in combination with Figure 4 As shown, the base is a trapezoidal structure with an arc-shaped upper bottom edge, which can be adapted to the outer wall of the test piece; more specifically, the support base 8 is arranged between adjacent clamping members 5 and is used to support the test piece 2, so that the test piece 2 and its adjacent clamping members 5 can simulate at least part of the convergence section of the thrust chamber 3.

[0074] In addition, the present application also includes a coolant input component for introducing coolant into the first cooling channel and a coolant output component for outputting coolant from the first channel.

[0075] Specifically, the coolant input component includes an input pipe, a mass flow meter arranged on the input pipe, a thermocouple arranged on the input pipe, and a pressure gauge arranged on the input pipe; further, the mass flow meter is used to measure the mass flow rate of the coolant input to the test piece 2, the thermocouple is used to measure the temperature of the coolant input to the test piece 2, and the pressure gauge is used to measure the pressure of the coolant input to the test piece 2.

[0076] Specifically, the coolant output component includes an output pipe, a mass flow meter arranged on the output pipe, a thermocouple arranged on the output pipe, and a pressure gauge arranged on the output pipe; further, the mass flow meter is used to measure the mass flow rate of the coolant output by the test piece 2, the thermocouple is used to measure the temperature of the coolant output by the test piece 2, and the pressure gauge is used to measure the pressure of the coolant output by the test piece 2.

[0077] In summary, (1) compared with material tests and thermal engine fatigue panel tests, the test bench design for thrust chamber structure tests can better reflect the influence of the complete thrust chamber morphology and structure, so its test results can better reflect the damage of the doghouse. (2) It can reflect the structural response of the inner wall of the thrust chamber composed of materials, but its cost is much lower than the thermal test of the entire thrust chamber. (3) Since the test piece is small, its structural surface temperature, pressure, etc. can be monitored in real time by measuring instruments during the test, so the details of the damage evolution of the test piece (panel) can be seen, helping researchers to gain a deeper understanding of the damage mechanism and build a more predictive material structure damage model.

[0078] Example 2

[0079] Combine Figure 5 and Figure 6 As shown, an embodiment of the present application also provides a thermomechanical fatigue testing method for a rocket thrust chamber material structure.

[0080] A method for thermomechanical fatigue testing of a rocket thrust chamber material structure comprises the following steps:

[0081] Preparation step 100: fix the test piece 2 with the clamping member 17; arrange the infrared thermometer 9 and the electromagnetic coil 10 above the test piece 2 with intervals and both facing the test piece 2; fill the sealed shell 1 with protective gas.

[0082] Specifically, step 101: the material of the clamping member 5 and the inner wall 13 of the test piece (where the inner side of the inner wall 13 is the gas side 15 and the outer side of the inner wall 13 is the atmospheric pressure side 16) is selected as oxygen-free copper OFHC; the total number of the first cooling channel 4 and the second cooling channel 7 is set to 300; the thickness of the inner wall 13 of the thrust chamber 3 is set to t in0.94mm; Setting the thickness of the outer wall 14 of the thrust chamber 3 t out The thickness w2 of the rib 18 formed between adjacent first cooling channels 4 (the parameters of this rib 18 and the rib 18 formed between adjacent second cooling channels 7) is set to 1.0 mm; the height h of the first cooling channel 4 (second cooling channel 7) is set to 6 mm; the width w1 of the first cooling channel 4 (second cooling channel 7) is set to 1.1 mm;

[0083] Step 102: Fix the test piece 2 on the support base 8; then use the clamping member 5 to clamp both ends of the test piece to simulate the partial convergence section structure of the thrust chamber 3; connect one end of the first cooling channel 4 to the input pipe 11, and the other end to the output channel;

[0084] Step 103: Arrange the infrared thermometer 9 and the electromagnetic coil 10 above the test object 2 with an interval and both facing the test object 2;

[0085] Step 104 : Nitrogen gas, which can be used as a protective gas, is filled into the sealed shell 1 to minimize the influence of water vapor in the environment on the test piece 2 .

[0086] Pre-cooling step 200: turning on the infrared thermometer 9 and introducing coolant into the first cooling channel 4;

[0087] Specifically, the introduction step 201 is to introduce the coolant into the first cooling channel 4 through the input pipe 11, and use the mass flow meter, thermocouple and pressure gauge to measure the mass flow of the coolant input to the test piece 2. Temperature T in and pressure P in ;

[0088] Cooling step 202: coolant flows through the first cooling channel 4 of the test piece 2 to pre-cool the test piece 2;

[0089] Step 203 of exporting: The coolant flowing through the first cooling channel 4 flows out through the output pipe 12, and the mass flow rate of the coolant output to the test piece 2 is measured by using a mass flow meter, a thermocouple and a pressure gauge. Temperature T out and pressure P out .

[0090] Harmless treatment step 204: the coolant flows out of the test piece 2 through the output pipe 12 and is harmlessly treated; further, a stop valve is provided on the output pipe 12 to control the pressure of the fluid flowing out to the atmospheric pressure side 16.

[0091] It is worth noting that the infrared thermometer 9 is turned on at the beginning of the pre-cooling step. The infrared thermometer 9 measures and reflects the temperature changes around the heating area of ​​the test piece 2 throughout the test, which serves as a temperature reference for the central area.

[0092] Thermal test step 300: After pre-cooling for 3 seconds, the heating component is turned on to heat the test piece 2 at a power of 50-80 kW and heat the test piece 2 from the original temperature to the target temperature, so as to achieve a high temperature environment simulating the actual situation.

[0093] Post-cooling step 400: After maintaining the target temperature for 500 seconds, the heating component is turned off and the temperature of the test piece 2 is cooled to the original temperature over 30 seconds. During the test, it was found that the temperature of the test piece 2 after cooling may be slightly higher than the original temperature;

[0094] Relaxation step 500 : Stop feeding the coolant into the first cooling channel 4 for 1000 seconds until the surface temperature of the test piece 2 becomes stable, thus completing one test cycle of the test piece 2 .

[0095] It is worth noting that the transition time between the preparation step 100, the pre-cooling step 200, the thermal test step 300, the post-cooling step 400 and the relaxation step 500 is 1 s.

[0096] Cycle step 600: repeat the preparation step, pre-cooling step, hot test step, post-cooling step and relaxation step N times in sequence, and observe the surface changes of the test piece 2.

[0097] Specifically, by following the aforementioned steps for cyclic testing, "doghouse damage" will eventually occur on the inner wall 13 of the test piece 2. After a cycle is complete, the life testing apparatus for the rocket thrust chamber 3 material can be closed, the sealed shell 1 opened, and the surface morphology of the test piece 2 observed. The cyclic thermal loading can then be continued until the test is complete.

[0098] In summary, the present application simulates a part of the convergence section of the thrust chamber 3 to replace the entire thrust chamber 3, uses a heating component to heat the workpiece to be tested to meet the high temperature conditions under the action of real gas, and uses safe and controllable liquid nitrogen (the coolant can also be liquid hydrogen or liquid methane) to cool the workpiece to simulate the thermal protection measures of the real situation. The heating process of the single-sided electromagnetic coil 10 can reflect the changes in the surface temperature of the workpiece to be tested.

[0099] Compared with the existing technology, it can measure various responses of the structure of the test piece 2, but it does not require the entire thrust chamber 3 to be tested in the ignition state after assembly, so the test cost is greatly reduced. At the same time, it can provide more comprehensive temperature and deformation measurements, reflecting the state evolution of the test piece 2 during the entire test process, so it is better helpful for studying its failure mechanism.

[0100] In addition, the surface temperature and pressure of the test piece 2 can be monitored in real time by measuring instruments, so the details of the failure evolution of the test piece 2 can be seen, helping researchers to gain a deeper understanding of the failure mechanism and build a more predictive material structure failure model.

[0101] Finally, it should be noted that 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 or all of the technical features therein. These modifications or replacements do not deviate the essence of the corresponding technical solutions from the scope of the technical solutions of the embodiments of the present application.

Claims

1. A thermomechanical fatigue testing device for rocket thrust chamber material structure, characterized in that: It comprises a sealed shell, a clamping member, a temperature measuring component and a heating component arranged in the sealed shell; The clamping member has at least two clamping parts, with a clamping space formed between adjacent clamping parts, the clamping space being used to clamp the test piece; the clamping parts and the test piece have the same curved surface, so that the adjacent clamping parts and the test piece clamped between the adjacent clamping parts can simulate at least a portion of the convergent section of the thrust chamber; The heating component faces the test piece and is used to heat the test piece; The temperature measuring assembly faces the test piece and is used to detect the temperature of the test piece. A first cooling channel for introducing a coolant is provided on the side wall of the test piece along its extending direction.

2. The thermomechanical fatigue testing device for rocket thrust chamber material structure according to claim 1, characterized in that: The clamping member includes a clamping piece and an end seat; There are multiple clamping members, and the clamping space is formed between adjacent clamping members; The end seat is arranged at an end of the clamping member away from the object to be tested.

3. The thermo-mechanical fatigue testing device for rocket thrust chamber material structure according to claim 2, characterized in that: A second cooling channel is provided on the side wall of the clamping member along its extending direction; the second cooling channel is used to simulate the cooling channel of the thrust chamber.

4. The thermo-mechanical fatigue testing device for rocket thrust chamber material structure according to claim 2, characterized in that: The clamping member further includes a support base; The support base is disposed between adjacent clamping members and is used to support the test piece, so that the test piece and its adjacent clamping members can simulate at least a portion of the convergent section of the thrust chamber.

5. The thermo-mechanical fatigue testing device for rocket thrust chamber material structure according to claim 1, characterized in that: The temperature measurement component is an infrared thermometer.

6. The thermo-mechanical fatigue testing device for rocket thrust chamber material structure according to claim 1, characterized in that: The heating assembly includes an electromagnetic coil, which can perform single-sided inductive heating on the surface of the test piece.

7. A method for thermomechanical fatigue testing of a rocket thrust chamber material structure, based on the thermomechanical fatigue testing device for a rocket thrust chamber material structure according to claim 5; characterized in that: The thermomechanical fatigue testing method for rocket thrust chamber material structures includes the following steps: Preparation steps: fix the test piece with a clamping member; arrange the infrared thermometer and the electromagnetic coil above the test piece with intervals and both facing the test piece; fill the sealed shell with protective gas; Pre-cooling step: turn on the infrared thermometer and introduce coolant into the first cooling channel; Thermal test step: After precooling for a first preset time, the heating component is turned on to heat the test piece and allow the test piece to reach the target temperature from the original temperature; Post-cooling step: After maintaining the target temperature for a second preset time, the heating component is turned off, and the temperature of the test piece is reduced to the original temperature after a third preset time; Relaxation step: stop supplying the coolant to the first cooling channel, and wait for a fourth preset time until the surface temperature of the test piece tends to be stable, thereby completing one test cycle of the test piece.

8. The thermomechanical fatigue testing method for rocket thrust chamber material structure according to claim 7, characterized in that: The precooling step further comprises the following steps: Introducing step: introducing coolant into the first cooling channel through the input pipe, and using a mass flow meter, a thermocouple, and a pressure gauge to respectively measure the mass flow, temperature, and pressure of the coolant input into the test piece; Cooling step: a coolant flows through a first cooling channel of the test piece to pre-cool the test piece; Output step: The coolant flowing through the first pipe flows out through the output pipe, and the mass flow rate, temperature and pressure of the coolant output to the test piece are measured respectively by using a mass flow meter, a thermocouple and a pressure gauge.

9. The thermomechanical fatigue testing method for rocket thrust chamber material structure according to claim 7, characterized in that: The relaxation step further includes the following steps: Cycle steps: Repeat the preparation step, pre-cooling step, hot test step, post-cooling step and relaxation step N times in sequence, and observe the surface changes of the test piece.

10. The thermomechanical fatigue testing method for rocket thrust chamber material structure according to claim 7, characterized in that: The coolant is liquid nitrogen.

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