A high-temperature impact-resistant electric heating sheet for spacecraft, and its preparation method and use

Abstract

The present invention discloses a high-temperature shock-resistant electric heater for spacecraft, relating to the technical field of electric heating elements. The heater comprises a three-layer composite structure, from top to bottom, comprising a composite film (thermal insulation layer) formed by in-situ chemical bonding of aerogel and polyimide (PI), and a polyimide composite film (thermal conductive insulation layer) composed of a double-layer Fe-Cr-Al / Ni-Cr alloy and vertically oriented carboxyl-functionalized carbon nanotubes. The heater exhibits high thermal conductivity, wide temperature stability, rapid thermal response, and high-temperature shock resistance. The preparation method includes preparing the thermal insulation layer, integrating the alloy heating layer with the thermal conductive insulation layer, strengthening the imidization reaction, and performing plasma-assisted heat treatment. This heater can be used for thermal management in extreme temperature environments, such as spacecraft propulsion systems or precision electronic equipment. It solves the problems of high-temperature stratification and response hysteresis associated with conventional heaters, thereby improving thermal control reliability.

Description

Technical Field

[0001] The present invention relates to the technical field of electric heating elements, in particular to a high-temperature impact-resistant electric heating sheet for spacecraft, a preparation method thereof, and uses thereof. Background Art

[0002] In the aerospace field, electric heaters are core components that ensure the reliable operation of spacecraft in extreme temperature environments. Their performance directly affects the stability of precision instruments, propulsion systems, and energy equipment. In space environments, spacecraft face drastic temperature fluctuations (such as from a deep cold shadow zone of -150°C to a high-temperature operating scene). Traditional electric heaters are difficult to meet these requirements due to material and structural defects:

[0003] (1) Interlayer bonding failure: Traditional electric heaters use physical bonding layers such as epoxy adhesives, which have an interface shear strength of less than 8 MPa. They are prone to delamination at temperatures above 300°C or under thermal shock, resulting in failure of the heating function. For example, when acrylic pressure-sensitive adhesive is used as a connecting layer, due to its low thermal conductivity, heating to above 300°C in milliseconds will cause carbonization and ablation of the interface, threatening the thermal management safety of the spacecraft.

[0004] (2) Imbalance between thermal conductivity and thermal insulation performance: The material properties of existing thermal conductive layers (such as graphene films) and thermal insulation layers (such as ceramic fibers) conflict, resulting in heat loss or local overheating (temperature difference > 8°C), making it impossible to achieve precise temperature control. For example, in deep space exploration, traditional systems are difficult to quickly de-ice due to thermal response hysteresis (> 50ms), which may cause risks such as frosting of optical devices and blockage of propellant lines.

[0005] (3) Insufficient temperature adaptability: The operating temperature range of most traditional electric heaters is limited to -50-300°C. The polymer matrix is ​​easily carbonized at high temperatures, and the material is brittle and fails in low temperature environments. For example, in extremely low temperature scenarios such as -180°C on the moon or in the polar regions of Mars, traditional electric heaters cannot work stably, which restricts the expansion of deep space exploration missions.

[0006] In view of this, this application is hereby filed. Summary of the Invention

[0007] The purpose of the present invention is to provide a high-temperature impact-resistant electric heating plate for spacecraft to solve the problems mentioned in the above background technology.

[0008] In order to solve the above technical problems, the present invention provides a high-temperature impact-resistant electric heating plate for spacecraft, wherein the electric heating plate has a three-layer composite structure, which includes, from top to bottom, a thermal insulation layer, an alloy heating layer, and a thermal conductive insulation layer;

[0009] The thermal insulation layer is a composite film formed by in-situ chemical bonding of aerogel and polyimide PI, with a gradient pore structure inside, a porosity of 85%-95%, and a thermal conductivity of ≤0.025W / (m·K);

[0010] The alloy heating layer is a double-layer alloy, the two layers are Fe-Cr-Al and Ni-Cr, and adopts a bionic fractal alloy circuit design with a resistivity range of 1-10Ω·cm;

[0011] The heat-conducting insulating layer is a polyimide composite film in which carboxyl functionalized carbon nanotubes are vertically oriented and arranged, and has a thickness of 10-50 μm.

[0012] Furthermore, the thermal insulation layer forms Si-O-Si covalent bonds through the mediation of siloxane coupling agent KH-550, realizing in-situ chemical bonding between aerogel and PI, with a thickness of ≤50 μm, a bending radius of ≤5 mm, and can withstand ≥1200°C.

[0013] Furthermore, the bionic fractal alloy circuit is one of a Hilbert curve and a tree-like fractal structure, with a thermal response time of <30ms and a thickness of 5-20μm.

[0014] Furthermore, the orientation degree of the carboxyl functionalized carbon nanotubes in the polyimide matrix is ​​≥90%, and they are prepared by electric field induction or template method, and have a longitudinal thermal conductivity of ≥45 W / (m·K) and a transverse resistivity of ≥10 15 Ω·cm.

[0015] A method for preparing a high-temperature shock-resistant electric heating sheet for a spacecraft comprises the following steps:

[0016] (1) Preparation of thermal insulation layer: Gradient spraying of silica aerogel precursor solution (such as ethyl orthosilicate ethanol solution) on the surface of PI film, adding KH-550 coupling agent, hot pressing at 200 °C and 10 MPa pressure for 20–30 min to form an aerogel-PI composite film;

[0017] (2) Integration of alloy heating layer and thermal conductive insulation layer: Fractal alloy circuits are prepared on a temporary substrate by laser direct writing technology, transferred to the surface of the vertically oriented PI-CNT film, and connected by resistance welding process;

[0018] (3) Imidization reaction enhancement: Polyamic acid (PAA) solution is coated on the surface of the aerogel-PI composite layer, covering the alloy heating layer and the PI-CNT thermal conductive layer, and the aerogel-PI composite layer is heated by dynamic humidity control (80% RH → 50% RH) and infrared radiation focused heating (50–100 W / cm 2 ), achieving imidization conversion rate>98%;

[0019] (4) Plasma-assisted heat treatment: In an argon plasma (50W, 13.56MHz) environment, the temperature was raised from 80℃ to 250℃ at a rate of 5℃ / min, so that the interlayer bonding energy was ≥0.8J / m 2 .

[0020] Furthermore, the preparation accuracy of the alloy heating layer in step (2) is ≤10 μm, and the wire welding contact resistance is ≤0.1 mΩ.

[0021] Furthermore, the spot diameter of the infrared radiation focused heating in step (3) is ≤2 mm, thereby achieving rapid imidization in a local area.

[0022] A high-temperature impact-resistant electric heating plate for spacecraft is used for thermal management of extreme temperature environments of spacecraft propulsion systems, new energy vehicle power batteries or precision electronic equipment. The operating temperature range of the electric heating plate is -196°C to 400°C. Furthermore, the operating temperature range of the electric heating plate is 300°C to 400°C.

[0023] Compared with the prior art, the present invention has the following beneficial effects:

[0024] 1. The interlayer bonding strength is significantly improved to solve the problem of high-temperature delamination failure:

[0025] Traditional electric heating sheets use a physical adhesive layer (such as epoxy glue) to connect the functional layers. The interfacial shear strength is usually less than 8MPa and is prone to delamination under high temperature (>300℃) or thermal shock. The present invention combines aerogel-PI in-situ chemical bonding (Si-O-Si covalent bond mediated by siloxane coupling agent) with plasma-assisted heat treatment (the bonding energy is increased to 0.8J / m 2 ), so that the interlayer shear strength reaches more than 15MPa, and there is still no interface delamination after 1000 times of 400℃ thermal shock cycle, which completely solves the problem of high-temperature delamination failure.

[0026] 2. Synergistic optimization of thermal conductivity and thermal insulation performance to achieve precise temperature control:

[0027] In existing technologies, thermally conductive layers (such as graphene membranes) and thermally insulating layers (such as ceramic fibers) are incompatible due to conflicting material properties, leading to heat loss or localized overheating (temperature differences >8°C). This new approach utilizes a composite design of a vertically oriented CNT-PI thermal conductive layer (with a longitudinal thermal conductivity of 45 W / (m·K)) and a gradient-pore aerogel-PI insulating layer (with a thermal conductivity of 0.025 W / (m·K)). This allows heat to be directed to the target area while suppressing lateral diffusion. The measured heating uniformity achieved a temperature difference of <2°C (tested on new energy vehicle battery modules), a 75% improvement over traditional solutions.

[0028] 3. The thermal response speed is improved dramatically to meet transient heating requirements:

[0029] Traditional serpentine alloy heating circuits typically have thermal response times exceeding 50ms due to their single path and uneven current density distribution. This invention utilizes biomimetic fractal alloy circuits (e.g., Hilbert curves) to increase the effective heating area by 30%, optimize the current path, and combine this with the rapid heat diffusion of the highly thermally conductive PI-CNT layer to shorten the thermal response time to 28ms (as tested with an infrared thermal imager), meeting the 40ms transient heating requirements of spacecraft thruster igniters.

[0030] 4. Significantly enhanced stability in extreme temperature ranges, expanding application scenarios:

[0031] Existing electric heaters are mostly limited to an operating temperature range of -50-300°C, and are prone to material degradation (such as carbonization of the polymer matrix) at high temperatures. This invention utilizes a high-temperature-resistant aerogel-PI composite layer (stable to 1200°C) and a double-layer alloy circuit (Fe-Cr-Al / Ni-Cr double-layer alloy). This allows the electric heater to maintain a resistance change of less than 2% (four-probe test) after 1000 thermal shock cycles within the -196-400°C range. This makes it suitable for use in extreme environments, from liquid nitrogen environments (-196°C) to high-temperature spacecraft components (400°C). BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Figure 1 The present invention is a flow chart of a method for preparing a high-temperature impact-resistant electric heating plate for spacecraft.

[0033] Figure 2 Temperature response curve after power on (a) Example 1; (b) Comparative Example 1

[0034] Figure 3 Temperature distribution diagram of the stable stage (a) Example 1; (b) Comparative Example 1

[0035] Figure 4 Material deformation diagram after 1000 minutes of power supply (a) Example 1; (b) Comparative Example 1 DETAILED DESCRIPTION

[0036] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0037] See also Figure 1 The present invention provides a technical solution: a high-temperature impact-resistant electric heating plate for spacecraft, wherein the electric heating plate is a three-layer composite structure, which includes, from top to bottom, a thermal insulation layer, an alloy heating layer, and a thermal conductive insulation layer;

[0038] The thermal insulation layer is a composite film formed by in-situ chemical bonding of aerogel and polyimide (PI), with a gradient pore structure inside, a porosity of 85%-95%, a thickness of ≤50μm, a thermal conductivity of ≤0.025W / (m·K), a bending radius of ≤5mm, and can withstand high temperatures of ≥1200°C;

[0039] The alloy heating layer is a double-layer Fe-Cr-Al / Ni-Cr alloy, adopts a bionic fractal alloy circuit design, has a resistivity range of 1-10Ω·cm, a thermal response time of <30ms, and a thickness of 5–20μm;

[0040] The Fe-Cr-Al alloy may be a 0Cr25Al5 alloy, wherein the Ni-Cr alloy comprises 65-85% Ni, 12-18% Cr, 0.5-2.5% Al, 1.0-3.0% Mo, and the balance Fe. Preferably, the Al / Mo ratio is between 0.4 and 1.2.

[0041] Compared with other thermally conductive alloys, the advantage of the Fe-Cr-Al / Ni-Cr double-layer alloy used in this application is that it can overcome the Cr depletion and performance degradation that occurs in Ni-Cr alloys during long-term use.

[0042] The thermally conductive insulating layer is a polyimide composite film with carboxyl functionalized carbon nanotubes arranged vertically, with a longitudinal thermal conductivity of ≥45W / (m·K) and a transverse resistivity of ≥10 15 Ω·cm, thickness is 10-50μm.

[0043] The thermal insulation layer forms Si-O-Si covalent bonds through the mediation of a siloxane coupling agent (KH-550), thereby achieving in-situ chemical bonding between the aerogel and the PI.

[0044] The bionic fractal alloy circuit is a type of Hilbert curve or tree-like fractal structure, and its effective heating area is increased by more than 30% compared with the traditional serpentine circuit.

[0045] The carboxyl functionalized carbon nanotubes have an orientation degree of ≥90% in a polyimide matrix and are prepared by electric field induction or a template method.

[0046] A method for preparing a high-temperature shock-resistant electric heating sheet for a spacecraft comprises the following steps:

[0047] (1) Preparation of thermal insulation layer: Gradient spraying of silica aerogel precursor solution (such as ethyl orthosilicate ethanol solution) on the surface of PI film, adding KH-550 coupling agent, hot pressing at 200 °C and 10 MPa pressure for 20–30 min to form an aerogel-PI composite film;

[0048] (2) Integration of alloy heating layer and thermal insulation layer: Fractal alloy circuits are prepared on a temporary substrate by laser direct writing technology, transferred to the surface of the vertically oriented PI-CNT film, and connected by resistance welding process; the preparation accuracy of the fractal alloy circuit is ≤10μm, and the contact resistance of the wire welding is ≤0.1mΩ;

[0049] (3) Imidization reaction enhancement: Polyamic acid (PAA) solution is coated on the surface of the aerogel-PI composite layer, covering the alloy heating layer and the PI-CNT thermal conductive layer, and the aerogel-PI composite layer is heated by dynamic humidity control (80% RH → 50% RH) and infrared radiation focused heating (50–100 W / cm 2 ), achieving an imidization conversion rate of >98%; the spot diameter of the infrared radiation focused heating is ≤2mm, achieving rapid imidization in a local area;

[0050] (4) Plasma-assisted heat treatment: In an argon plasma (50W, 13.56MHz) environment, the temperature was raised from 80℃ to 250℃ at a rate of 5℃ / min, so that the interlayer bonding energy was ≥0.8J / m 2 .

[0051] A high-temperature impact-resistant electric heating sheet for spacecraft is used for thermal management of extreme temperature environments of spacecraft propulsion systems, new energy vehicle power batteries or precision electronic equipment, with an operating temperature range of -196°C to 400°C.

[0052] To make the purpose, technical solutions and advantages of the present invention clearer, the present invention will be described in detail below with reference to specific embodiments. It should be noted that, unless there is a conflict, the embodiments and features in the embodiments of the present application can be combined with each other.

[0053] Example 1:

[0054] 1 Preparation of thermal insulation layer

[0055] Aerogel precursor solution: Tetraethyl orthosilicate (TEOS), ethanol, and deionized water were mixed in a volume ratio of 5:3:2, 0.5% hydrochloric acid was added to adjust the pH to 2, and the mixture was stirred at room temperature for 2 hours to form a silica sol.

[0056] Gradient spraying: Spray the silica sol evenly on the surface of 50μm thick PI film, the spraying thickness is controlled to 30μm (wet film), and sprinkle KH-550 coupling agent powder (dosage 5mg / cm 2 ).

[0057] Hot pressing process: Hot pressing at 200°C and 10 MPa pressure for 25 minutes to form an aerogel-PI composite film (thickness 45 μm, porosity 90%, thermal conductivity 0.02 W / (m·K)).

[0058] 2 Alloy heating layer integrated with thermal insulation layer

[0059] Fractal alloy circuits: Hilbert-shaped Fe-Cr-Al / Ni-Cr bilayer alloy circuits (20 μm width, 10 μm thickness, 5 Ω·cm resistivity) were fabricated on a quartz substrate using laser direct writing. These circuits were then transferred to a vertically oriented PI-CNT film (15% carboxylated CNT content, longitudinal thermal conductivity 50 W / (m·K)) by hot pressing. The Fe-Cr-Al alloy consisted of 0Cr25Al5, and the Ni-Cr alloy consisted of 75% Ni, 15% Cr, 2.0% Al, 2.0% Mo, and the balance Fe.

[0060] Wire welding: Silver-based solder (melting point 280°C) was used to connect copper wires with a diameter of 0.1 mm by resistance welding, with a contact resistance of 0.08 mΩ.

[0061] 3. Enhanced imidization reaction

[0062] PAA solution coating: PAA solution with a solid content of 15% (thickness 20 μm) is coated on the surface of the aerogel-PI composite layer, covering the alloy heating layer and the PI-CNT thermal conductive layer.

[0063] Co-processing: First, place it in an 80% RH humidity environment for 30 minutes, then heat it with infrared focus (spot diameter 1.5mm, power 80W / cm 2 ) The temperature was raised to 350°C and maintained for 1 hour, and the imidization conversion rate was 99%.

[0064] 4 Plasma assisted heat treatment

[0065] Processing conditions: Ar flow rate 20sccm, plasma power 50W, heating from 80℃ to 250℃ at a rate of 5℃ / min, holding for 30min, interlayer bonding energy reached 0.9J / m 2 .

[0066] Performance testing:

[0067] Interlaminar shear strength: 16.2MPa (ASTMD1002 standard), no peeling after 1000 cycles of 400℃ thermal shock.

[0068] Heating uniformity: temperature difference 1.8°C (new energy vehicle battery module test, temperature range 25°C→60°C).

[0069] Thermal response time: 28ms (tested with infrared thermal imager, heating to 300°C).

[0070] Temperature stability: -196℃ liquid nitrogen immersion for 30min + 400℃ high temperature baking for 1 hour, resistance change rate is 1.5% (four-probe method).

[0071] Resistance change rate after 1000 cycles at high temperature: After 1000 cycles at 0-400°C, measure the resistance change rate, referring to GB / T 2423.22-2012.

[0072] Example 2: Optimization of the pore structure of the thermal insulation layer

[0073] The difference between Example 2 and Example 1 is that the thermal insulation layer is designed with a bottom layer (contact heating layer) having a porosity of 85% and a top layer having a porosity of 95%, which is achieved by spraying silica sols of different concentrations (sol solid content of the bottom layer is 10% and that of the top layer is 5%).

[0074] Other processes are the same as in Example 1.

[0075] Performance improvements:

[0076] Thermal conductivity: 0.018W / (m·K) (top layer has enhanced thermal insulation).

[0077] Bending radius: 3mm (increased flexibility), suitable for heating curved parts.

[0078] Example 3:

[0079] The difference between Example 3 and Example 1 is that an aerogel-PI composite film (thickness 45 μm, porosity 50%) is formed in the thermal insulation layer.

[0080] Example 4: Comparison of Alloy Circuit Materials

[0081] Example 4 differs from Example 1 in that the heating layer uses a Ni-Cr alloy (resistivity 8Ω·cm) and a tree-like fractal circuit design, increasing the effective heating area by 40%. The Ni-Cr alloy consists of 75% Ni, 15% Cr, 2.0% Al, 2.0% Mo, and the balance Fe.

[0082] Other processes are the same as in Example 1.

[0083] Example 5: Comparison of Alloy Circuit Materials

[0084] Example 5 differs from Example 1 in that the heating layer uses an Fe-Cr-Al alloy (resistivity 8Ω·cm) and a tree-like fractal structure, increasing the effective heating area by 40%. The Fe-Cr-Al alloy is 0Cr25Al5.

[0085] Other processes are the same as in Example 1.

[0086] Comparative Example 1: Traditional electric heating sheet (230 type flexible film electric heater)

[0087] Structure: Epoxy adhesive bonded PI insulation layer + serpentine Ni-Cr alloy circuit + aluminum foil thermal conductive layer.

[0088] Process: physical bonding, no plasma treatment.

[0089] As shown in Table 1 below:

[0090] In order to better illustrate the superiority of the present invention, a general-purpose aerospace electric heater was selected as a comparison item, and indoor tests and numerical simulation analyses were carried out. Figure 2 The results of an indoor thermal response test after power-on were obtained. The test item (inventive item) stabilized its temperature 31 seconds after power-on, reaching an average temperature of 304°C. The comparison item (general-purpose type) stabilized its temperature 42 seconds after power-on, reaching an average temperature of 261°C. The results show that the improved electric heater significantly improved its thermal response efficiency and stable temperature, with thermal response efficiency increasing by 35.5% and stable temperature increasing by 16.5%.

[0091] The finite element simulation software Comsol is used to simulate the temperature distribution of the invention and general electric heating plates after thermal stabilization. Figure 3 The figure shows the temperature distribution after stabilization of the inventive solution. The self-similarity of the Hilbert fractal function makes the spatial distribution of the heating filaments more uniform, significantly reducing the temperature difference between the heating lines and the empty areas. The inventive solution controls the temperature difference of the general item from 5°C to within 2°C. This reduced temperature difference also significantly reduces the warping and deformation caused by thermal expansion and contraction of the heat transfer and heating layers. Numerical simulation results show that the displacement of the inventive solution is reduced by 0.006mm compared to the traditional solution, and the warping phenomenon is reduced by 2.1%.

[0092] Figure 4 Figures 1 and 2 show material deformation after 1000 minutes of power-on (a) Example 1 and (b) Comparative Example 1, indicating that the deformation of Example 1 is significantly smaller than that of Comparative Example 1.

[0093] Table 1: Performance comparison of the present invention and comparative examples

[0094]

[0095] As can be seen, the effects of Example 2 and Example 1 are similar, but the effect of Example 3 is inferior, possibly due to the porosity. Different alloy materials were used in Examples 1, 4, and 5, resulting in Example 1 having better performance than Examples 4-5. Meanwhile, the performance of Comparative Example 1, representing a conventional electric heater, is significantly inferior and does not meet the requirements.

[0096] In summary, in the present invention:

[0097] (1) Using aerogel-PI in situ chemical bonding (Si-O-Si covalent bond mediated by siloxane coupling agent) and plasma assisted thermal treatment (binding energy increased to 0.8 J / m 2), so that the interlaminar shear strength reaches more than 15MPa, and there is still no interface delamination after 1000 cycles of 400℃ thermal shock.

[0098] (2) Through the composite design of a vertically oriented CNT-PI thermal conductive layer (longitudinal thermal conductivity 45W / (m·K)) and a gradient pore aerogel-PI thermal insulation layer (thermal conductivity 0.025W / (m·K)), heat is directed to the target area while suppressing lateral diffusion. The measured heating uniformity temperature difference is <2°C (new energy vehicle battery module test), which is 75% higher than the traditional solution.

[0099] (3) By adopting bionic fractal alloy circuits (such as Hilbert curves), the effective heating area is increased by 30% and the current path is optimized. Combined with the rapid thermal diffusion of the high thermal conductivity PI-CNT layer, the thermal response time is shortened to 28ms (infrared thermal imager test), meeting the 40ms transient heating requirement of the spacecraft thruster igniter.

[0100] (4) Through the high temperature resistant design of the aerogel-PI composite layer (1200℃ stability) and the double-layer alloy circuit (Fe-Cr-Al / Ni-Cr double layer), the resistance change rate of the electric heater is less than 2% after 1000 hot and cold shocks in the range of -196-400℃ (four-probe test), which is suitable for extreme scenarios from liquid nitrogen environment (-196℃) to high temperature components of spacecraft (400℃).

[0101] Finally, it should be noted that the above embodiments are merely illustrative of the technical solutions of the embodiments of the present invention, and are not intended to limit them. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they may still modify the technical solutions described in the aforementioned embodiments, or replace some of the technical features therein with equivalents; and such modifications or replacements do not deviate from the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present invention. Any changes or replacements that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be covered by the scope of protection of the present invention.

Claims

1. A high temperature impact resistant electric heating sheet for spacecraft, characterized by: The electric heating plate is a three-layer composite structure, which includes, from top to bottom: a thermal insulation layer, an alloy heating layer, and a heat conductive insulation layer; The thermal insulation layer is a composite film formed by in-situ chemical bonding of aerogel and polyimide PI, with a gradient pore structure inside, a porosity of 85%-95%, and a thermal conductivity of ≤0.025W / (m·K); The alloy heating layer is a double-layer alloy, the two layers are Fe-Cr-Al and Ni-Cr, and adopts a bionic fractal alloy circuit design with a resistivity range of 1-10Ω·cm; The heat-conducting insulating layer is a polyimide composite film in which carboxyl functionalized carbon nanotubes are vertically aligned, and has a thickness of 10-50 μm.

2. The high temperature impact resistant electric heater for spacecraft according to claim 1, characterized in that: The thermal insulation layer forms Si-O-Si covalent bonds through the mediation of siloxane coupling agent KH-550 to achieve in-situ chemical bonding between aerogel and PI, has a thickness of ≤50 μm, a bending radius of ≤5 mm, and can withstand temperatures of ≥1200°C.

3. The high temperature impact resistant electric heater for spacecraft according to claim 1, characterized in that: The bionic fractal alloy circuit is one of a Hilbert curve and a tree-like fractal structure, has a thermal response time of less than 30ms, and a thickness of 5-20 μm.

4. The high temperature impact resistant electric heater for spacecraft according to claim 1, characterized in that: The carboxyl functionalized carbon nanotubes have an orientation degree of ≥90% in the polyimide matrix and are prepared by electric field induction or template method. The longitudinal thermal conductivity is ≥45 W / (m·K) and the transverse resistivity is ≥10 15 Ω·cm,.

5. The method for preparing the electric heating sheet according to any one of claims 1 to 4, characterized in that: The following steps are involved: (1) Preparation of thermal insulation layer: Spray the silica aerogel precursor solution onto the surface of the PI film in a gradient manner, add KH-550 coupling agent, and hot press at 200°C and 10 MPa pressure for 20-30 minutes to form an aerogel-PI composite film; (2) Integration of alloy heating layer and thermal conductive insulation layer: Fractal alloy circuits are prepared on a temporary substrate by laser direct writing technology, transferred to the surface of the vertically oriented PI-CNT film, and connected by resistance welding process; (3) Enhancement of imidization reaction: Polyamic acid PAA solution is coated on the surface of the aerogel-PI composite layer, covering the alloy heating layer and the PI-CNT thermal conductive layer. Through dynamic humidity control and infrared radiation focused heating, the imidization conversion rate is achieved to be >98%; (4) Plasma-assisted heat treatment: In an argon plasma environment, the temperature is raised from 80°C to 250°C at a rate of 5°C / min to make the interlayer bonding energy ≥ 0.8J / m 2 .

6. The high temperature impact resistant electric heater for spacecraft according to claim 5, characterized in that: The preparation accuracy of the alloy heating layer in step (2) is ≤10 μm, and the wire welding contact resistance is ≤0.1 mΩ.

7. The high temperature impact resistant electric heater for spacecraft according to claim 5, characterized in that: The spot diameter of the infrared radiation focused heating in step (3) is ≤2 mm, so as to achieve rapid imidization in a local area.

8. Use of the electric heating sheet according to any one of claims 1 to 4, characterized in that: Used for thermal management of spacecraft propulsion systems or precision electronic equipment in extreme temperature environments.

9. The use according to claim 8, characterized in that: The operating temperature range of the electric heating plate is -196℃-400℃.

10. The use according to claim 8, characterized in that: The operating temperature range of the electric heating plate is 300℃-400℃.

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