CN-122009468-A - Heat protection structure of multi-layer hypersonic aircraft and preparation method thereof

CN122009468ACN 122009468 ACN122009468 ACN 122009468ACN-122009468-A

Abstract

A heat protection structure of a multi-layer hypersonic aircraft and a preparation method thereof relate to the technical field of hypersonic aircraft heat protection and comprise a C/SiC composite material layer, a vacuum buffer layer, a white calcium carbonate layer, a solar cell module and a thermoelectric generation sheet which are sequentially arranged from outside to inside, wherein the C/SiC composite material layer directly resists extreme high temperature and ablation through five layers of collaborative design; the vacuum buffer layer blocks solid heat conduction and reduces heat transfer inwards; the integrated design of blocking-reflecting-converting improves the heat protection efficiency by more than 40%, can control the internal structure temperature within the safety range of less than or equal to 100 ℃ in the extreme environment of 2000-3000 ℃, and realizes the comprehensive performance balance of high temperature resistance, ablation resistance, low heat conduction and high toughness.

Inventors

XU SHINAN
LIU XUDONG
HUANG TIANHAO
DONG JINGPING
LI MINGHONG
LI YUFENG
ZHANG QIANG

Assignees

四川铁道职业学院

Dates

Publication Date: 20260512
Application Date: 20260415

Claims (10)

1. A heat protection structure of a multi-layer hypersonic aircraft comprises a C/SiC composite material layer, a vacuum buffer layer, a white calcium carbonate layer, a solar cell module and a thermoelectric generation sheet which are sequentially arranged from outside to inside, and is characterized in that a square frame is fixed outside the vacuum buffer layer, the vacuum buffer layer comprises a support piece and a buffer spring which are arranged between the C/SiC composite material layer and the white calcium carbonate layer, the support piece is of an I-shaped structure and comprises a cylindrical piece and extension rings at the upper end and the lower end of the cylindrical piece, the extension rings of the support piece are respectively bonded and fixed with the inner wall of the C/SiC composite material layer and the outer wall of the white calcium carbonate layer through high-temperature-resistant ceramic adhesive, bulges are respectively fixed at the upper part and the lower part of the cylindrical piece, the buffer spring is sleeved outside the cylindrical piece, two ends of the buffer spring respectively penetrate through perforations on the bulges and are in interference fit with the perforations, the solar cell module comprises a plurality of solar cell panels, each solar cell panel is connected in series or in parallel in a square array structure, the solar cell panel is fixedly connected with the white calcium carbonate layer through titanium alloy supports at the two ends, and the contact part of the solar cell panel and the thermoelectric generation panel is coated with the thermoelectric generation panel.
2. The heat protection structure of the multi-layer hypersonic aircraft according to claim 1 is characterized in that high temperature resistant ceramic adhesive is Al 2 O 3 -based adhesive, the temperature resistance is more than or equal to 1600 ℃, high temperature alloy is Inconel718, the solar cell panel is a GaAs-based high temperature solar cell, the power density is more than or equal to 50W/m < 2 >, the working temperature is-50 ℃ to 200 ℃, the photoelectric conversion efficiency is more than or equal to 25%, the thermoelectric generation sheet is Bi 2 Te 3 -based thermoelectric material, the power density is more than or equal to 10W/m < 2 >, the Seebeck coefficient is more than or equal to 200 mu V/K, the thickness is 3mm to 5mm, and the heat conductivity coefficient of the heat conducting silica gel is more than or equal to 1.5W/(m.k).
3. The heat protection structure of the multi-layer hypersonic aircraft according to claim 1 is characterized in that a titanium alloy bracket is of an L-shaped structure and has a thickness of 1mm-2mm, mo alloy bolts for fixing a solar cell panel and a white calcium carbonate layer are respectively arranged on the titanium alloy bracket in a penetrating manner, and the temperature resistance of the Mo alloy bolts is more than or equal to 600 ℃.
4. The heat protection structure of the multi-layer hypersonic aircraft according to claim 1 is characterized in that a preformed hole connector is arranged on the square frame, a ceramic sealing connector and a molecular pump connector are respectively fixed in the preformed hole connector, a lead wire on the solar panel passes through the ceramic sealing connector and then is connected with an outgoing wire on the thermoelectric generation sheet in parallel after passing through the square frame, and the molecular pump connector is connected with an external molecular pump vacuum port.
5. A method for producing a thermal protection structure for a multi-layer hypersonic aircraft according to any one of claims 1 to 4, characterized in that it comprises in particular the following steps: S1, preparing a C/SiC composite material layer by adopting an improved CVI method; S2, processing a support piece and a buffer spring of the vacuum buffer layer, and after assembling the support piece and the buffer spring, bonding one end of the support piece to the inner wall of the C/SiC composite material layer through high-temperature resistant ceramic glue; s3, sintering the white calcium carbonate layer, and adhering and fixing the other end of the supporting piece with the white calcium carbonate layer; S4, fixedly connecting the solar cell panel with the white calcium carbonate layer through the titanium alloy bracket and the Mo alloy bolt to complete the whole assembly; s5, carrying out vacuum treatment and integral sealing on the assembled heat protection structure: The interface of the vacuum buffer layer is connected with a molecular pump, the molecular pump is firstly pumped to the rough vacuum of 10Pa, then the molecular pump is started to 5 multiplied by 10 -4 Pa, the pressure is maintained for 30 minutes, and the pressure change is less than or equal to 1Pa; Closing the vacuum valve, sealing the interface with copper gasket, coating heat-conducting silica gel on the contact surface of the solar cell panel and the white calcium carbonate layer and the contact surface of the solar cell panel and the thermoelectric generation sheet, filling with heat-conducting silica gel, coating C/SiC cloth belt on the surface, and binding and fixing with ceramic wire.
6. The method for preparing the heat protection structure of the multi-layer hypersonic flight vehicle according to claim 5, wherein the specific steps for preparing the C/SiC composite material layer in the step S1 are as follows: s1.1, selecting T700-grade carbon fiber, and preparing a carbon fiber preform by adopting a 3D four-way braiding process, wherein the fiber volume fraction is 40% -45%; S1.2, placing the carbon fiber preform into a horizontal CVI reaction furnace, heating to 900 ℃, and introducing inert gas Ar for purifying for 2 hours; s1.3, introducing reaction gases SiCl 4 and CH 4 , wherein the gas flow ratio is 3:1, the flow rate of SiCl 4 is 80sccm-100sccm, the flow rate of CH 4 is 25sccm-30sccm, and the pressure in the furnace is controlled at 0.3MPa, so that vapor deposition is completed; S1.4, performing pitch impregnation and pyrolysis densification operation on the carbon fiber preform subjected to the vapor deposition in the previous step; S1.5, stopping depositing for 4.5 weeks through a rotating tool at a rotating speed of 5r/min, wherein the thickness of the formed C/SiC composite material layer is 2mm-3mm, the grain size of the SiC matrix is 5 mu m-10 mu m, and the temperature is resistant to 1800 ℃ to 2200 ℃.
7. The method for preparing the heat protection structure of the multi-layer hypersonic aircraft according to claim 5, wherein the specific operation of performing pitch impregnation in the step S1.4 is that the carbon fiber preform subjected to vapor deposition is placed into a special high-pressure impregnation kettle, the inside of the special high-pressure impregnation kettle is firstly vacuumized until the vacuum degree is less than or equal to 5Pa, and the inside of the preform is kept for 30min to remove residual air, then molten mesophase pitch with the softening point of 200-220 ℃ is injected into the special high-pressure impregnation kettle, the temperature is raised to 250 ℃ and the pressure of 0.8-1.0 MPa is applied, and the impregnation time is 2-3 h, so that the pitch is fully impregnated into micropores of the carbon fiber preform; The pyrolysis densification operation comprises the steps of transferring the carbon fiber preform impregnated with asphalt to an inert atmosphere sintering furnace, introducing argon with the flow rate of 50-60 sccm as a protective atmosphere, heating to 1000-1200 ℃ at the speed of 5-min, preserving heat for 2 hours to complete pyrolysis, and repeating the soaking-pyrolysis process for 1-2 times according to the porosity requirement.
8. The method for preparing the heat protection structure of the multi-layer hypersonic flight vehicle according to claim 5, wherein the specific processing steps of the vacuum buffer layer in the step S2 are as follows: S2.1, preparing a supporting piece by adopting the CVI method which is the same as that of the previous step, wherein the supporting piece is made of a C/SiC composite material, the machining size is 5mm-8mm in diameter and 10mm-15mm in height, the surface is subjected to sand blasting treatment, the roughness Ra=5 mu m, and protrusions with perforations are machined at two ends of a cylindrical piece of the supporting piece; S2.2, processing an Inconel718 alloy wire into a spiral buffer spring, wherein the free length is 8mm-12mm, the number of turns is 5, the elastic coefficient is 12N/m, and the compression amount is less than or equal to 5mm; S2.3, marking a mounting point of a support piece on the inner wall of the C/SiC composite material layer 1, coating one end of the support piece with high-temperature-resistant ceramic glue, bonding the support piece at the marked mounting point, and curing for 3 hours in a 250 ℃ oven; S2.4, sleeving the buffer spring outside the support piece cylindrical piece, enabling two end parts of the buffer spring to penetrate through holes on protrusions on the upper side and the lower side of the cylindrical piece, enabling the square frame to be a C/SiC thin plate, enabling the upper end face and the lower end face of the square frame to be respectively welded and connected with the lower face of the C/SiC composite material layer and the upper face of the white calcium carbonate layer, enabling a reserved hole connector to be formed in the square frame, and fixing a ceramic sealing connector and a molecular pump connector in the reserved hole connector.
9. The method for preparing the heat protection structure of the multi-layer hypersonic flight vehicle according to claim 5, wherein the specific preparation steps in the step S3 are as follows: S3.1, firstly, drying the calcium carbonate powder in a vacuum drying oven at 120 ℃ for 4 hours, and sieving the dried powder with a 200-mesh sieve; S3.2, filling a graphite mould, applying 5MPa pressure for prepressing forming, putting into a box furnace, sintering for 3 hours at 300-400 ℃ in air atmosphere for forming, naturally cooling to room temperature, demoulding, and carrying out surface grinding treatment until Ra is less than or equal to 1.6 mu m, thickness is 0.5-2 mm, and reflectivity is more than or equal to 85%; And S3.3, coating high-temperature-resistant ceramic glue on the other end of the support piece, bonding with the white calcium carbonate layer, and curing for 3 hours in a 250 ℃ oven.
10. The method for preparing the heat protection structure of the multi-layer hypersonic aircraft according to claim 5 is characterized in that the step S4 of installing the solar panel, the white calcium carbonate layer and the thermoelectric generation sheet is as follows: S4.1, connecting solar panels in series/parallel to form a square array, adhering titanium alloy brackets on two sides of the solar panels, and fixing the solar panels with a white calcium carbonate layer through Mo alloy bolts on the titanium alloy brackets; S4.2, attaching the thermoelectric generation sheet to the lower surface of the solar cell panel, coating heat-conducting silica gel on the contact surface, applying pressure of 0.5MPa to maintain the pressure for 1 hour, adhering the graphite heat-conducting sheet on the cold surface, connecting an outgoing line on the thermoelectric generation sheet with a circuit of the solar cell panel in parallel, and connecting the total output end to an aircraft power supply system through a high-temperature-resistant cable.

Description

Heat protection structure of multi-layer hypersonic aircraft and preparation method thereof Technical Field The invention relates to the technical field of hypersonic aircraft heat protection, in particular to a heat protection structure of a multi-layer hypersonic aircraft and a preparation method thereof. Background The hypersonic aircraft refers to an aircraft with cruising speed exceeding Mach 5, can reach a battlefield outside 5000 km within 1 hour, has the advantages of good concealment, strong burst prevention capability, high combat efficiency, wide combat space and the like, can realize strategic targets such as accurate striking, rapid transportation, remote reconnaissance and the like, and is a new sky strategy high point. However, when in flight, the hypersonic aircraft surface and high-speed air flow are subjected to intense friction to generate pneumatic heating, so that the surface temperature is rapidly increased, thermal stress is generated on the internal structure of the aircraft, the strength is reduced, and the structural integrity is damaged; the heat-resistant composite material is characterized in that the heat-resistant composite material is mainly composed of a C/C composite material, a C/C composite material and a ceramic material, wherein the C/C composite material is used for radiating heat in a single mode, the C/C composite material is used for conducting heat in a high thermal conductivity mode, but is poor in oxidation resistance, the structure is provided with the problems of serious heat accumulation, low energy utilization rate and insufficient structural stability, 2, the existing composite material is used for enabling a single coating to fall off due to thermal expansion or vibration and impact at high temperature to enable the structure to fail, 3, the existing heat protection technology is used for guaranteeing high temperature resistance, a thick-wall single material is usually adopted, the thickness of an ultrahigh-temperature ceramic layer is 5-10mm, the structural weight is large, the range and maneuverability of an aircraft are influenced, 4, the existing ultrahigh-temperature composite material is also provided with an obvious short plate, the C/C composite material is required to be protected by an extra coating, the C/SiC composite material is low in thermal conductivity, the heat is easy to be accumulated, UHTC ceramic is low in toughness and easy to be fragile, and therefore, the heat protection structure of the multilayer type high-speed aircraft and the preparation method thereof become a basic requirement for technicians in the field. Disclosure of Invention The invention aims to overcome the defects of the prior art and provides a heat protection structure of a multi-layer hypersonic aircraft and a preparation method thereof. The technical scheme adopted by the invention is as follows: The utility model provides a multilayer hypersonic aircraft's thermal protection structure, including the C/SiC combined material layer that sets gradually from outside to interior, vacuum buffer layer, white calcium carbonate layer, solar module and thermoelectric generation piece, be fixed with square frame in the outside of vacuum buffer layer, the vacuum buffer layer includes the support piece that sets up between C/SiC combined material layer and white calcium carbonate layer and buffers the spring, the support piece is "worker" font structure, including the cylinder piece and the extension ring at the upper and lower both ends of cylinder piece, the extension ring of support piece is fixed with the inner wall of C/SiC combined material layer respectively through high temperature resistant ceramic adhesive bonding, the outer wall of white calcium carbonate layer, upper portion and lower part at the cylinder piece are all fixed with the arch, buffer spring cover is established in the outside of cylinder piece, and buffer spring's both ends pass the perforation on the arch respectively, with perforation interference fit, the solar module includes the polylith solar cell panel, the polylith solar panel is square array structure in series or parallelly connected for square array structure, every solar panel is connected with white calcium carbonate layer through titanium alloy support at both ends, solar panel and the contact with the solar panel, the solar panel and the thermoelectric generation piece is connected with the solar panel of silica gel in contact department that the solar panel and the thermoelectric generation piece all has the contact with the silica gel. The heat protection structure of the multi-layer hypersonic aircraft is characterized in that the high-temperature resistant ceramic adhesive is Al 2O3 -based adhesive, the temperature resistance is more than or equal to 1600 ℃, the high-temperature alloy is Inconel718, the solar panel is a GaAs-based high-temperature solar cell, the power density is more than or equal to 50W/m < 2 >, the working tempera