Ceramic Matrix Composites (CMCs)

Ceramic matrix composites (CMCs) reinforced with nanocarbon have attracted significant interest due to their potential to enhance mechanical, thermal, and electrical properties. Although the investigation of carbon-based materials such as graphene and carbon nanotubes as additives for advanced ceramics has been widespread, the utilization of metal-organic framework (MOF)-derived nanocarbons in CMCs remains largely unexplored. We extended our previous proof-of-concept investigations by demonstrating the effectiveness of a different type of MOF-derived carbon as a reinforcing phase in an alternative ceramic matrix. We employed spark plasma sintering (SPS) to consolidate yttria-stabilized zirconia (YSZ) and zeolitic imidazolate framework (ZIF-67) powder blends at 1300 °C and a uniaxial pressure of 50 MPa. YSZ serves as the ceramic matrix, whereas ZIF-67 serves as the nanocarbon source. The composite exhibits a highly significant improvement in fracture toughness with an increase of up to 13% compared to that of the YSZ monolith. The formation of ZIF-derived nanocarbon interlayers is responsible for the observed enhancement in ductility, which can be attributed to their ability to facilitate energy dissipation during crack propagation and inhibit grain growth. Furthermore, the room-temperature electrical conductivity of the sintered samples demonstrates a substantial improvement, primarily due to the in situ formation of nanocarbon-based fillers, reaching an impressive 27 S/m with 10 wt % ZIF-67 content. Based on the results, it can be inferred that the incorporation of in situ MOF-derived nanocarbons into CMCs leads to a substantial improvement in both the mechanical and electrical properties.

Long fibre reinforced ceramics with active fillers and a modified intra-matrix bond based on the LPI process

There exists today considerable interest in developing continuous fiber-reinforced ceramic matrix composites (CMC) that can operate as hot-section components in advanced gas turbine engines. The objective of this paper is to present simple analytical and empirical models for predicting the effects of time and temperature on CMC tensile rupture under various composite and engine conditions. These models are based on the average rupture behavior measured to air for oxide and SiC-based fibers of current technical interest. For example, assuming a cracked matrix and Larson-Miller rupture curves for single fibers, it is shown that model predictions agree quite well with high-temperature stress-rupture data for SiC/SiC CMC. Rupture models, yet to be validated, are also presented for three other relevant conditions: (a) SiC fibers become oxidatively bonded to each other in a cracked CMC, (b) applied CMC stresses are low enough to avoid matrix cracking, and (c) Si-based CMC are subjected to surface recession in high temperature combustion gases. The practical implications of the modeling results are discussed, particularly in regard to the optimum fibers and matrices for CMC engine applications and the thermostructural capability of SiC/SiC CMC in comparison to nickel-based superalloys, monolithic ceramics, and oxide/oxide CMC.

Ceramic matrix composites (CMCs), in particular oxide-based systems, are of interest for use in combustion turbines. While uncoated oxide CMCs have significant hurdles to implementation in gas turbines, the Siemens hybrid oxide CMC system is able to overcome these challenges. These hybrid oxide CMCs provide distinct advantages over the current non-oxide based systems. The benefits of hybrid oxide-based systems for advanced gas turbines will be discussed. Material system developments will be discussed including those completed by a Siemens Power Generation led team in a recent NIST (National Institute of Standards and Technology) sponsored program to prove the concept of advanced hybrid oxide-oxide CMCs for gas turbine engines. The program fabricated a full scale outer combustor liner that was installed in a Solar Centaur 50S engine at a commercial end user site. In November 2006, this hybrid oxide CMC outer combustor liner met the target life goal of 25,000 hrs with 25,404 hrs of field test experience. The final hurdle for design of hybrid oxide CMC components is the ability to accurately analytically predict behavior. Methods and approaches to address this challenge are discussed as well.

Ceramic matrix composites (CMCs) have lower density and a higher service temperature limit than nickel based alloys which have been used for turbine components of aircraft engines. These properties of CMCs have the potential to reduce the weight of turbine components and improve turbine thermal efficiency with a higher turbine inlet temperature (TIT). One of the technical issues of the CMC turbine vane is a relatively lower impact resistance than nickel based alloy turbine vanes. There are various previous works about impact resistance of CMCs, but there is little work that assumed actual engine conditions. The objective of this work was to verify the resistance of SiC/SiC CMC turbine vane to the impact phenomena that occur in the actual aircraft engine. The field damage survey was conducted on actual metal turbine vanes of commercial engines overhauled in IHI. The survey made it clear that the typical damage was less-than-0.127-mm-dent at the leading edge. In addition, the dropped weight impact test using the actual turbine airfoil which is made from a nickel based alloy was conducted at ambient temperature. The amount of energy required to make the dent of a certain size that was observed in actual metal turbine vanes was estimated. Then, the dropped weight impact test using the CMC test piece with a leading edge shape was conducted at the impact energy estimated by the metal turbine airfoil. The results showed that the failure mode of the CMC test piece was local damage with dents of a certain size and not a catastrophic failure mode. From this work, the damage to be assumed on CMC vane in actual aircraft engines was identified. As a future task, the effect of the damage to the fatigue capability of CMC turbine vanes needs to be investigated.

Ceramic Matrix Composites (CMCs) are materials in which ceramics are reinforced with fibers or other materials, providing enhanced functionalities, such as damage tolerance, in addition to their inherent heat resistance. This study investigates a manufacturing method for continuous alumina-mullite fiber-reinforced alumina CMCs using a commercial Fused Deposition Modeling (FDM) 3D printer, demonstrating a flexible and practical route for continuous fiber-reinforced CMC fabrication. The process utilizes a pre-impregnated filament, which serves as the fundamental component for depositing both the ceramic matrix and continuous reinforcement. The 3D printing filament was fabricated by melting a mixture of alumina, serving as the matrix material, and a thermoplastic resin to impregnate the alumina-mullite mixed fibers. Filaments with a diameter of 0.4 mm exhibited fewer cracks compared to those with a diameter of 0.6 mm. Moreover, a filament with fewer voids was produced when the heating temperature during impregnation was 230 °C. Using the fabricated filament, simple three-layered rectangular test specimens, with a length of 110 mm and a width of 10 mm, were successfully printed. A tensile strength of approximately 250 MPa was achieved along the fiber direction. The observed fracture behavior indicates that improving the interlaminar strength would likely further enhance the tensile strength. The impact of this research is introducing a new manufacturing method that will expand the design freedom and accessibility of complex-shaped CMCs. This pioneering achievement provides a foundational baseline for researchers and engineers to transition 3D printed continuous fiber-reinforced CMCs from laboratory-scale prototypes to structural components.

Towards physical properties tailoring of carbon nanotubes-reinforced ceramic matrix composites

With the rapid development of advanced manufacturing technologies and the growing demand for high-performance materials, zirconia ceramic-epoxy resin composites have demonstrated enormous potential in fields such as aerospace (e.g., lightweight and high-temperature resistant components) and biomedicine (e.g., bone implants). However, homogeneous ceramic structures fabricated via traditional sintering processes struggle to balance high strength and lightweight properties, while brittle crack propagation remains a prominent issue. Based on vat photopolymerization additive manufacturing technology, this study utilizes Triply Periodic Minimal Surfaces (TPMS) as the design foundation to construct Gyroid-type porous ceramic matrices through topological optimization. Epoxy resin is introduced via vacuum impregnation to fill the pores, leveraging the energy absorption characteristics of the polymer phase to inhibit crack propagation. Experimental results indicate that the zirconia ceramic-epoxy resin composite achieves a compressive strength of 100 MPa, representing a 250% improvement compared to the pure ceramic matrix. Both energy absorption (EA) and specific energy absorption (SEA) are enhanced by approximately 70 times, with a relative density controlled at 25%. This research provides a novel approach for the customized design of high-performance ceramic matrix composites (CMCs) and holds significant application value in aerospace engine thermal insulation components and personalized orthopedic implants.

Silicon-based ceramics are attractive materials for use in gas turbine engine hot sections due to their high temperature mechanical and physical properties as well as lower density than metals. The advantages of utilizing ceramic hot section components include weight reduction, and improved efficiency as well as enhanced power output and lower emissions as a result of reducing or eliminating cooling. Potential gas turbine ceramic components for industrial, commercial and/or military high temperature turbine applications include combustor liners, vanes, rotors, and shrouds. These components require materials that can withstand high temperatures and pressures for long duration under steam-rich environments. For Navy applications, ceramic hot section components have the potential to increase the operation range. The amount of weight reduced by utilizing a lighter gas turbine can be used to increase fuel storage capacity while a more efficient gas turbine consumes less fuel. Both improvements enable a longer operation range for Navy ships and aircraft. Ceramic hot section components will also be beneficial to the Navy's Growth Joint Strike Fighter (JSF) and VAATE (Versatile Affordable Advanced Turbine Engines) initiatives in terms of reduced weight, cooling air savings, and capability/cost index (CCI). For DOE applications, ceramic hot section components provide an avenue to achieve low emissions while improving efficiency. Combustors made of ceramic material can withstand higher wall temperatures and require less cooling air. Ability of the ceramics to withstand high temperatures enables novel combustor designs that have reduced NO{sub x}, smoke and CO levels. In the turbine section, ceramic vanes and blades do not require sophisticated cooling schemes currently used for metal components. The saved cooling air could be used to further improve efficiency and power output. The objectives of this contract were to develop technologies critical for ceramic hot section components for gas turbine engines. Significant technical progress has been made towards maturation of the EBC and CMC technologies for incorporation into gas turbine engine hot-section. Promising EBC candidates for longer life and/or higher temperature applications relative to current state of the art BSAS-based EBCs have been identified. These next generation coating systems have been scaled-up from coupons to components and are currently being field tested in Solar Centaur 50S engine. CMC combustor liners were designed, fabricated and tested in a FT8 sector rig to demonstrate the benefits of a high temperature material system. Pretest predictions made through the use of perfectly stirred reactor models showed a 2-3x benefit in CO emissions for CMC versus metallic liners. The sector-rig test validated the pretest predictions with >2x benefit in CO at the same NOx levels at various load conditions. The CMC liners also survived several trip shut downs thereby validating the CMC design methodology. Significant technical progress has been made towards incorporation of ceramic matrix composites (CMC) and environmental barrier coatings (EBC) technologies into gas turbine engine hot-section. The second phase of the program focused on the demonstration of a reverse flow annular CMC combustor. This has included overcoming the challenges of design and fabrication of CMCs into 'complex' shapes; developing processing to apply EBCs to 'engine hardware'; testing of an advanced combustor enabled by CMCs in a PW206 rig; and the validation of performance benefits against a metal baseline. The rig test validated many of the pretest predictions with a 40-50% reduction in pattern factor compared to the baseline and reductions in NOx levels at maximum power conditions. The next steps are to develop an understanding of the life limiting mechanisms in EBC and CMC materials, developing a design system for EBC coated CMCs and durability testing in an engine environment.

We are conducting the development of ceramic matrix composites (CMC) and components made of CMC for a 100 kW automotive ceramic gas turbine (CGT) as shown in Fig.1. When compared to monolithic ceramics (MC), CMC that we have developed demonstrate superior strength characteristics in terms of resistance to particle impact and thermal shock. We have conducted evaluation tests on the strength of CMC components in which MC such as silicon nitride and silicon carbide were used as a reference for comparison with CMC in the same testing process as employed for components made of MC such as silicon nitride and silicon carbide. It was confirmed that actual components made of CMC realized approximately the same strength as the test pieces. Furthermore, some CMC components have already passed screening tests that evaluated the strength of the components. It was therefore confirmed that the potential exists for the possibility of testing these components in high temperature assembly tests and engine tests.

Surface treatment effects on ceramic matrix composites: Case of a thermal sprayed alumina coating on SiC composites

Fatigue of unitized polymer/ceramic matrix composites with 2D and 3D fiber architecture at elevated temperature

An approach to predictions of failure in ceramics and of failure damage in Ceramic Matrix Composites (CMC) is discussed. For ceramics, particular emphasis was placed upon the populations of fracture inducing flaws. The flaw strength parameters, that are used to infer the failure of various loading geometries were estimated in view of fracture origins identified by scanning electron microscopy, using analytical as well as computerized methods. The proposed approach for CMC considers the microcomposite geometry. The Weibull model as well as a more fundamental model based upon the elemental strength concept were used.

A progressive damage model for ceramic matrix mini-composites with thick interphases

The objective of this effort is to predict ceramic matrix composites (CMC) interlaminar Mode II Crack Growth Resistance (CGR), and the design of ASTM test specimen. Currently, there are a number of test standards and American Society for Testing and Materials (ASTM) for CMC’s at both ambient and elevated temperatures; however, there are no standardized test methods for determination of interlaminar shear (Mode II) fracture toughness in CMC’s. Although research work exists on interlaminar Mode II fracture toughness of CMC’s, the test methods applied showed definite drawbacks and limitations. Delamination Crack Growth (CGR) tests of CMC Mode II may exhibit zig-zag pattern, wavy cracks, fiber bridging, and premature specimen failure under bending load. The experimental parameters that may contribute to the difficulty can be summarized as specimen width and thickness, interface coating thickness, mixed mode failure evolution, and interlaminar defects. Modes II crack growth resistances, GII, were analytically and numerically determined at ambient temperature using end notched flexure (ENF) and the end-loaded split (ELS). Finite Element (FE) based. Multi-scale progressive failure analysis (MS-PFA) a combined Micro-mechanical damage and fracture mechanics Virtual Crack Closure Technique (VCCT) algorithms. Modeling of melt-infiltrated SiC/SiC CMC of ENF specimen (Laminate: with initial crack length was accomplished using a MS-PFA and VCCT approach. Test data were compared with MS-PFA prediction: a) Force vs. Crack Opening Displacement; and b) Mode II crack tip energy release rate vs. crack extension length for both edge and center line due to formation of Micro Crack Density Contribution, Crack Tip Stiffness Reduction; and c) zig-zag crack growth behavior (adhesive/cohesive). Next the ASTM Standard Proposed linear SGR equation was developed based on interpretation compliance technique from both MS-PFA Analysis and Test.