| Polyimide-based composite materials are widely used in aerospace structural parts because of their better heat resistance than other resin-based composite materials.However,with the rapid development of aerospace technology,the heat resistance of composite materials and high temperature.The lower service life puts forward higher requirements.This article is aimed at the polyimide carbon fiber(CFPI)composite material used in aviation structural parts,and the high thermal conductivity graphite film(HTCGF)is compounded on the surface through the hot-press co-curing process,and the structure thermal protection mechanism and material damage under high temperature environment are explored.The assessment and thermal oxygen aging mechanism,and the establishment of a life prediction model,laid the foundation for the subsequent application of composite materials at high temperatures.In this paper,the carbon fiber polyimide prepreg was prepared by the solution dipping method;the prepreg was layered according to the process requirements,and the high thermal conductivity graphite film was laid on the top layer;the high thermal conductivity graphite film reinforced by the autoclave was prepared Carbon fiber resin-based composite material;cut it into corresponding sizes according to the experimental requirements;conduct thermal oxygen aging tests at different temperatures and different heating rates through a customized muffle furnace;finally,test and analyze the structure and performance of the samples.1)The in-plane thermal conductivity of the graphite film is about 200 times higher than in the thickness direction.In the process of heat conduction,the heat is first transferred and diffused in the plane,and a better thermal protection effect can be achieved in the thickness direction;2)Mechanical performance test The results show that the bending properties and interlaminar shear properties are not monotonically increasing with the increase of heating time,but when the heating time is shorter,the mechanical properties show an upward trend,and gradually decrease with the increase of time;3)Combined with the analysis of the macroscopic morphology and the infrared thermal image,it can be found that when the heating rate is faster and the processing temperature is higher,because the small molecule gas inside the composite material cannot overflow well,the bulging phenomenon will occur at the interface of the graphite film,and the composite The thermal protection structure of the material will be damaged to varying degrees,causing the resin matrix to withstand a temperature higher than its working temperature,resulting in degradation of the material and a decrease in mechanical properties;4)Infrared and TG test analysis results show that the PI resin is degraded The temperature range is mainly between 450?-700?,combined with microscopic morphology analysis,it can be found that the debonding phenomenon will occur at the interface of carbon fiber and resin during the degradation process,and some small molecule particles will be produced at the same time,combined with TG-IR The results show that the main products of the degradation process are CO,CO2,H2O and other small molecular gases.According to the thermal analysis kinetics and the TG test results of the composite material,using the Arrhenius model and the Flynn-Wall-Ozawa method,combined with the thermogravimetric analysis(TG)data at different heating rates,the apparent activation energy E is calculated to be 165.75 KJ·Mol-1,obtain the pyrolysis dynamics equation of the material under high temperature service environment.Through a large number of experimental tests,the corresponding relationship between the heat source temperature,the film surface temperature and the substrate temperature is summarized.After adding a high thermal conductivity graphite film thermal protection structure,the temperature of the substrate can be reduced by 100-350? at a working temperature of 400-800?,after long time working at 650?,the temperature of the film surface is maintained at 500?,and the substrate temperature is stable at 315?,the material can maintain its excellent properties,which can explain the high temperature resistance of the material.A substantial increase.The corresponding relationship between the temperature of the heat source and the temperature of the substrate is brought into the above-mentioned life prediction model,and the life prediction equation is further obtained:t=1/A*ln(M/M0)exp(E/R(0.625 x+182.802)),thus It is predicted that the maximum use temperature of the composite material is 700?,and it can be used for about 15 h at this temperature,and it can be used for about 100 hours at a heat source temperature of 650?.According to the predicted results,it has been verified by multiple experiments.The heat source is set to heat up to 650? at a heating rate of 2?/min.After 48 h of thermal oxygen aging treatment,the mass loss rate of the material is only 0.5%,which is in line with the life prediction results.And material usage requirements.It proves that the life prediction model obtained in this paper has relatively high reliability and can lay a certain foundation for the subsequent application of thermal protection structure in aerospace.In short,this thesis provides a new idea to improve the high temperature resistance of resin matrix composites.Through the research of this thesis,we can grasp the key data such as the conduction heat transfer law and high temperature resistance of high thermal conductivity graphite film resin matrix composites.The temperature distribution law under thermal shock and the relationship between mechanical properties and heat,establish a life prediction system suitable for aviation composite materials,combine with relevant experimental verification and analysis,improve the reliability of the use of materials,and provide certain theoretical support for the practical application of composite materials. |
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