Monolithic ceramic is a kind of brittle material. The properties of the material will be greatly reduced by microdefects, which limit the practical application of ceramics in many fields. However, the inherent brittleness of ceramic materials can be improved by continuous or discontinuous ceramic fiber or carbon fiber reinforcement, namely, ceramic-matrix composites (CMCs). This dispersed second phase can improve the fracture toughness of ceramic materials. The main mechanism is that the crack bridging effect in the process of fracture can make the matrix materials connect with each other, disperse the fracture energy by the way of fiber debonding, and fiber pulling out to prevent the fracture of the material [1, 2].
Compared with the superalloy, fiber-reinforced CMCs can withstand higher temperature, reduce cooling air flow, and improve the turbine efficiency. The density of fiber-reinforced CMCs is 2.0ā2.5 g/cm3, which is only 1/4ā1/3 of superalloy. CMCs have already been applied to aeroengine combustion chambers, nozzle flaps, turbine vanes, and blades. For example, the CMC nozzle flaps and seals manufactured by SNECMA have already been used for more than 10 years in the M88 and M53 aeroengines. The CMC tail nozzle designed by SAFRAN Group of France passed the commercial flight certification of European Union Aviation Safety Agency (EASA) and completed its first commercial flight on the CFM56-5B aeroengine on 16 June 2015. National Aeronautics and Space Administration (NASA) has prepared and tested the CMC turbine guide vanes and turbine blade disc components in the Ultra-Efficient Engine Technology (UEET) project. General Electric (GE) tested the CMC combustor and high-pressure turbine components in the ground test of GEnx aeroengine. The CMC turbine blades were successfully tested on the F414 engine, which are planned to be used in GE90 series aeroengines. The engine weight is expected to be reduced by 455 kg, accounting for about 6% of the total weight of GE90-115 aeroengine. The LEAP (Leading Edge Aviation Propulsion) series aeroengine developed by CFM company adopts CMC components. The LEAP-1A, 1B, and 1C aeroengine provides power for Airbus A320, Boeing 737MAX, and COMAC C919, respectively.
Compared with polymer matrix composites (PMCs), CMCs have some similarities, including anisotropy, braided structure, high strength/high modulus fibers, manufacturing process sensitivity, and diversity, but there are also differences, such as high operation temperature (>500 Ā°C), diversity of material constituents (i.e. oxide matrix, nonoxidized matrix, carbide matrix, silicon nitride matrix, carbon matrix, etc.) and processing methods (i.e. reaction bonding [RB], hot pressing sintering [HPS], precursor infiltration and pyrolysis [PIP], reactive melt infiltration [RMI], chemical vapor infiltration [CVI], slurry infiltration and hot pressing [SIPH], CVI-PIP, CVI-RMI, and PIP-HP), low matrix failure strain, complex degradation/damage/failure mechanisms at elevated temperature, difficult connection of structures in high-temperature environment, and high requirement of nondestructive testing and repair technology.
To ensure the reliability and safety of their use in aircraft and aeroengine structures, it is necessary to investigate the tensile, fatigue, stress rupture, and vibration behavior of CMCs at elevated temperature.
1.1 Tensile Behavior of CMCs at Elevated Temperature
The stressāstrain curve of CMCs under tensile load appears obviously nonlinear. The tensile stressāstrain curve can be divided into three regions, i.e. elastic region, nonlinear region, and secondary linear region before failure. In region I, there is no damage in the material during initial loading, and the tensile stressāstrain curve is linear. With the increase of stress, microcracks appear in the matrix-rich area or matrix defects. The initial matrix cracking stress is defined as Ļmc. These microcracks can only be detected by means of acoustic emission (AE), microscopic observation of specimen surface, and temperature measurement of sample surface. When the stress reaches the proportional limit stress, the accumulation of matrix cracks makes the stressāstrain curve deflect, and the stressāstrain curve is nonlinear, which marks the beginning of region II. In region II, the matrix cracks propagate along the vertical load direction while the number of matrix cracks increases. When the cracks extend to the fiber/matrix interface, the cracks deflect along the fiber/matrix interface, and debonding occurs at the fiber/matrix interface. With the increase of stress, when the slip zones of adjacent matrix cracks overlap each other, the matrix cracks reach saturation. The saturated stress of matrix cracks is defined as Ļsat. When the matrix crack is saturated, the region III of the stressāstrain curve starts, and the external load is mainly borne by the fiber. The tangent modulus of the stressāstrain curve is about VfEf (Vf is the volume content of the fiber and Ef is the elastic modulus of the fiber). With the increase of the stress, some fibers fail, and the failed fibers continue to carry through the shear stress at the fiber/matrix interface. When the fibers broken fraction reaches the critical value, the composite fracture occurs.
The tensile stressāstrain behavior reflects the strength of the composite material to resist the damage of external tensile loading. The tensile properties (i.e. proportional limit stress, matrix crack spacing, tensile strength, and fracture strain) can be obtained from the tensile stressāstrain curves and can be used for component design [3ā5]. Jia et al. [6] investigated the relationship between the interphase and tensile strength of SiC fiber monofilament. The tensile strength of the SiC fiber monofilament decreases with the increasing coating layers. The SiC fibers with single boron nitride (BN) coating have the high monofilament strength retention of about 70%, 42.1% with two BN coatings, and 3...
