纤维增强陶瓷基复合材料的发展及应用

陶瓷基复合材料（CMC）是在陶瓷中引入第二相材料,使之增强,增韧的多相材料.法、美、德、日等发达国家20多年来积极开展研究,目前陶瓷基复合材料已在航空航天发动机、航天热防护系统、原子能等领域已经得到应用。     

关于先进陶瓷材料结构功能一体化的思考与实践

先进陶瓷是为满足现代科学技术发展的需要而开发的各种新型陶瓷的总称,可分为结构陶瓷和功能陶瓷两大类。结构陶瓷是利用先进陶瓷的高强度、高硬度、耐高温、耐腐蚀、化学稳定好等特性,被广泛用作各种高温结构材料、耐磨材料、耐腐蚀材料,如涡轮发动机叶片、火箭喷嘴、陶瓷轴承、切削刀具、磨料磨具、汽车尾气净化器等。     

超高温抗氧化复合材料研究进展

超高温抗氧化材料技术是制约新一代可重复使用液体火箭发动机技术发展的一项重要关键技术．为此国外近年来大力发展超高温(2200～3000℃)抗氧化材料技术。综述了国外新一代可重复使用液体火箭发动机用超高温抗氧化材料的研究进展,主要包括难熔金属及合金、难熔陶瓷材料、难熔碳化物、难熔硼化物和抗氧化C／C复合材料。最后指出了超高温抗氧化材料技术的研究方向。     

连续纤维增强陶瓷基复合材料国外应用研究进展

连续纤维增强陶瓷基复合材料(CFRCMCs)因具有高比强、高比模、高可靠性、耐高温等优异性能,已成为航空航天、军事、能源等领域理想的高温结构材料.主要介绍了国外CFRCMCs在热防护系统、火箭发动机和超燃冲压发动机上应用的研究进展.     

双组元液体动力环境下3D C/SiC复合材料喷管烧蚀性能

为明确C/SiC陶瓷基复合材料喷管在液体火箭发动机工作环境的烧蚀特性,采用先驱体浸渍-裂解（PIP）工艺制备得到3D C/SiC复合材料喷管,并对喷管进行多种工况环境下的地面热试车烧蚀考核。结果表明：制备得到的3D C/SiC复合材料喷管能够满足液体火箭发动机多种工况环境下抗烧蚀性能要求,喷管喉部线烧蚀率约为3.92×10^-4 mm/s;热试车烧蚀实验后喷管入口圆柱段、收敛段、喉部及扩张段外型面均残留有大量白色物质SiO₂,喉部局部出现烧蚀坑洞现象;C/SiC复合材料液体火箭发动机工作环境下的烧蚀机理为机械冲刷烧蚀和氧化烧蚀。     

陶瓷基复合材料或将革新航空发动机工作效率

正NASA正在寻求改变未来商用飞机的机遇,更有效的发动机是核心。为了实现未来飞机用上更好的发动机的目标,NASA的研究人员正在研究有前景的涡轮发动机部件高温材料。这些金属,被称为陶瓷基复合材料或CMC,更轻、更强,且能够给承受喷气发动机核心部件的极端高温环境下的受力要求。陶瓷基复合材料将替代目前在航空发动机中应用的镍基高温合金。总的来说,发动机工作温度越高,燃油效率更高。近年来,由于金属部件采用了热障涂层,发动机工作温     

连续纤维增强陶瓷基复合材料国外应用研究进展

连续纤维增强陶瓷基复合材料（CFRCMCs）因具有高比强、高比模、高可靠性、耐高温等优异性能，已成为航空航天、军事、能源等领域理想的高温结构材料。主要介绍了国外CFRCMCs在热防护系统、火箭发动机和超燃冲压发动机上应用的研究进展。     

固体火箭发动机喷管用C/C复合材料的研究进展

C/C复合材料已成为固体火箭发动机(SRM)最理想的喷管材料.综述了国内外SRM喷管用C/C复合材料的研究进展、应用状况以及烧蚀性能的研究状况,展望了今后C/C复合材料喷管的发展.     

聚合物先驱体溶液浸渍/裂解法制备3D C/SiC CMC的结构与性能

一、前言 连续纤维增强的SiC/SiC和C/SiC陶瓷基复合材料(CMC)由于具有高强、高模,低密度、良好的韧性,抗热震和良好的高温稳定性等一系列优点。目前国内外在高温结构材料方面正在竞相开发。除了一维和二维织物增强的CMC外,近年来由于三维编织技术的进步,3DCMC也开始得到研究。与1D,2D CMC相比,由于三维纤维交叉排列组成的整体网络结构,不仅可以显著改善其径向力学性能,提供复合材料的结构韧性机制,使其径向强度和韧性进一步提高,而且还可以仿形(nearnet shape)编织,从而使高温复合材料可直接成型为勿需机械加工的异形结构部件。目前制备3D CMC的方法主要有泥浆浸渍法和CVI法,泥浆浸渍法主要只用于玻璃基或玻璃陶瓷基,CVI法可用于制备各种氧化物和非氧化物陶瓷基,但设备复杂,工艺周期长。 聚合物先驱体转化法制备CMC由于具有工艺简单,成型温度低等优点,最近受到国内外重视但用此法制备3D CMC尚未见报道。 本文主要借鉴制备3D C/C复合材料的成功经验,探讨了采用聚合物先驱体溶液浸渍法制备3D C/SiC的方法,初步研究了3D C/SiC的结构与性能特点。     

离子火箭发动机真空试验设备的设计

根据离子火箭发动机试验的特殊要求,通过对真空室的尺寸和形状、离子火箭发动机的工作压力、真空室材料和主抽气真空泵选择的分析,得到满足离子火箭发动机试验要求的真空试验设备的设计。     

Near net-shape, ultra-high melting, recession-resistant ZrC/W-based rocket nozzle liners via the displacive compensation of porosity (DCP) method

Dense, near net-shaped ZrC/W-based composites have been fabricated at modest temperatures and at ambient pressure by a reactive infiltration process known as the Displacive Compensation of Porosity (DCP) method. Porous WC preforms with hourglass shapes (for rocket nozzle liners) were produced by gel casting, whereas simple bar-shaped preforms were produced by uniaxial pressing. The porous preforms were exposed to molten Zr₂Cu at 1200–1300°C and ambient pressure. The Zr₂Cu liquid rapidly infiltrated into the preforms and underwent a displacement reaction with the WC to yield a more voluminous mixture of solid products, ZrC and W. This displacement reaction-induced increase in internal solid volume filled the prior pore spaces of the preforms (displacive compensation of porosity) to yield dense, ZrC/W-based composites. Because the preforms remained rigid during reactive infiltration, the final composites retained the external shapes and dimensions of the starting preforms. A DCP-derived, ZrC/W-based nozzle insert was found to be resistant to the severe thermal shock and erosive conditions of a Pi-K rocket motor test. The DCP process enables dense, ceramic/refractory metal composites to be fabricated in complex and near net shapes without the need for high-temperature or high-pressure densification or for extensive machining (i.e., relatively expensive processing steps are avoided).     

某型号固体火箭发动机喷管型面设计与数值计算

根据某型号固体火箭发动机总体设计要求及限制条件,给出了喷管型面设计方法。采用计算流体动力学(CFD)方法,用单方程（Spalart-Allmaras）湍流模型,分析计算了不同入口压力下喷管内流场特性和壁面压力、温度分布等。根据流场计算结果求出了发动机瞬时推力,并与地面推力试验结果进行了对比。     

Numerical simulation of the dynamic thermostructural response of a composite rocket nozzle throat

The finite element method, in the form of the commercial finite element code ADINA, is used to investigate the dynamic thermostructural response of a composite rocket nozzle throat. ADINA’s thermoelastic analysis capability is validated by the comparison of its solution for the thermoelastic response of a thick, homogeneous, cylindrically orthotropic tube heated internally, to an analytical one. The spatially reinforced Carbon–Carbon nozzle throat examined here forms part of a low-erosion solid rocket motor nozzle model that is subjected to structural and thermal loading, with the effects of material ablation being neglected. An initial transient quasi-static thermostructural analysis is performed to determine the validity of the nozzle design, following which, an uncoupled dynamic thermostructural analysis of the nozzle’s throat and entrance section for the initial transient phase of the nozzle’s operation, is carried out. The results of this analysis are then compared to those of the equivalent transient quasi-static analysis to assess the degree of variance in either solution. It is found that the dynamic response oscillates about the quasi-static response in all cases, and that, in general, the variance in stress magnitudes between the two solution techniques is significant.     

固体火箭发动机喷管型面的研究进展

固体火箭发动机喷管型面设计直接关系到喷管效率和推力大小,是喷管设计中的重要研究课题。本文从喷管型面设计方法、型面参数优化和喷管流固热耦合分析等方面综述了国内外对固体火箭发动机喷管型面的研究进展。总结出了直接优化方法、特型喷管的设计方法、六次Bézier曲线、双三次样条曲线构造扩张段型线和B-Spline曲线和特征线方法等喷管型面的设计方法,并介绍了计算流体动力学(CFD)和随机优化方法在固体火箭发动机(SRM)设计优化中的运用。分析了固体火箭发动机喷管涉及到的流固热耦合问题,并结合文献介绍了经典CFD算法、CBS有限元算法和格子波尔兹曼在研究流固热耦合问题上的运用。     

梯度陶瓷喷嘴残余应力的有限元分析

为了提高喷嘴的抗冲蚀磨损能力,将梯度功能材料理论运用于喷嘴材料的设计中,改传统的均质喷嘴材料为非均质喷嘴材料,提出在梯度陶瓷喷嘴制备中将残余压应力引入喷嘴入口的设计目标.在组成分布指数一定的条件下,针对主要设计参数对梯度陶瓷喷嘴残余应力的影响进行有限元分析,探讨了梯度层厚度、临界梯度层材料组分差对SiC/(W,Ti)C单梯度陶瓷喷嘴残余热应力的影响规律,在组成分布指数取0.5时,优化SiC/(W,Ti)C梯度陶瓷喷嘴梯度层厚和临界梯度层材料组分差.结果表明,残余应力随梯度层厚h及临界梯度层SiC体积组分差的不同产生很大差异,合理设计梯度层厚h及临界梯度层SiC体积组分差可在喷嘴入口形成有效残余压应力,最佳梯度层厚为5mm,临界梯度层SiC组分差小于5%(体积分数).     

固体火箭发动机喷管阻尼数值计算方法及规律研究

从声学角度出发,参考稳态波衰减法,综合考虑喷管辐射损失及对流损失,建立固体火箭发动机喷管阻尼声能共振数值计算方法,通过与阻抗管法测量的试验结果对比验证仿真方法,并探究喷管喉径、监测点位置、声源强度及平均压力对喷管阻尼的影响,结果表明:数值计算结果与试验结果吻合较好,表明该研究建立的仿真方法有效;喷管阻尼随喉径的增大而增...     

Ablation,mechanical and thermal conductivity properties of vapor grown carbon fiber/phenolic matrix composites

The ablation, mechanical and thermal properties of vapor grown carbon fiber (VGCF) (Pyrograf III™ Applied Sciences, Inc.)/phenolic resin (SC-1008, Borden Chemical, Inc.) composites were evaluated to determine the potential of using this material in solid rocket motor nozzles. Composite specimens with varying VGCF loadings (30–50% wt.) including one sample with ex-rayon carbon fiber plies were prepared and exposed to a plasma torch for 20 s with a heat flux of 16.5 MW/m² at approximately 1650°C. Low erosion rates and little char formation were observed, confirming that these materials were promising for rocket motor nozzle materials. When fiber loadings increased, mechanical properties and ablative properties improved. The VGCF composites had low thermal conductivities (approximately 0.56 W/m-K) indicating they were good insulating materials. If a 65% fiber loading in VGCF composite could be achieved, then ablative properties are projected to be comparable to or better than the composite material currently used on the Space Shuttle Reusable Solid Rocket Motor (RSRM).     

Development of a Net Shape Manufacturing Method for Ventilated Brake Discs in Single Piece Design

Carbon fibre reinforced ceramic matrix composites (CMC), originally developed for lightweight heat shields of spacecraft, are used for high performance brake discs in sports cars from different manufacturers. In contrast to the CMC materials for space applications, based on woven fabrics and costly manufacturing methods, these low cost friction materials are produced by liquid silicon infiltration of porous Carbon/Carbon (C/C) preforms, based on short fibre reinforced CFRP green bodies manufactured via warm press technique. In this work, different manufacturing methods for ventilated CMC brake discs are compared to each other, and the development of a new technology for the manufacture of single piece C/C‐SiC brake discs in net shape technique is presented.     

Polymer derived ceramic matrix composites

Preceramic polymers offer a unique method to fabricate ceramic matrix composites (CMC). Relatively large and complex shapes were fabricated using a polysilazane polymer and silicon carbide based reinforcements of CG Nicalon™ and HI-nicalon™ fibers. This paper summarizes a raw material system and the fabrication process used to prepare two-dimensional cloth reinforced composites. Typical tensile, shear and compressive properties of CMCs prepared with the two types of reinforcements are presented. Although CG Nicalon reinforced composites exhibit good mechanical stability at moderate stress levels at 1100°C, HI-Nicalon reinforced composites show improved creep behavior at 1200°C.