wwceramics. org/ACT Auid oxides and the in-depth diffusion of oxygen is established and joining techniques under development. slowed down or stopped. An even more efficient strat- Similar materials could also be used for the shielding of egy consists in replacing the homogeneous SiC-matrix satellites against the impact of meteorites or foreign ob- itself by an engineered multilayered matrix based on a jects owing to the high toughness and hardness of these similar principle, referred to as a self healing matrix and materials that could be infiltrated by P-CVI at the laborator Another promising field of application is that of the scale.Here, the matrix is deposited as the repetition of hot structures of aerojet engines and related gas turbines a given sequence S comprising thin layers X acting as which are presently made of heavy, low melting point mechanical fuses(X being C, C B), BN, BN (Si), or any nickel-based superalloys that require complex coolin suitable fuse) and layers Y of species forming fluid ox- systems. Replacing superalloys by SiC-based composites ides, as mentioned above(Fig. 3b). Durability of the would permit to raise the gas temperature, suppress,or 1100@C has been reported for SiC fiber composites fab- ciency of the engine, and reduce both the weight anda) ricated with such self-healing matrices. Finally, noise/pollution level(Fig. 4b). However, it will proba ific multilayered coatings containing oxide layers bly take some time(these materials are still your as mullite or/and baryum strontium aluminosilicate, their fabrication costly) and be limited in a first step to BSAS) have been proposed and tested for SiC/SiC com- nonrotating parts, i. e,, the combustors and the after- posites exposed to wet oxidizing atmospheres to reduce burner parts such as the flaps of the exhaust nozzles. The the recession rate of the materials, e.g., in hot combu main concern here is durability, that should be of the tion gas rich in water vapor order of several thousands of hours. The outer(diver gent) Aaps of exhaust nozzles experience a temperature R applic that is relatively low(T<700oC). Hence, they can be fabricated with carbon fiber-reinforced SiC-matrix, with Space and Aeronautic Field a weight gain of 50%. Tested in Hight as early as 1989, they are now in volume production(M88 Snecn ma en- SiC-matrix composites are potential material can- gines of the Rafale fighter, Paris, France). 00, The didates for the fabrication of hot structures of spacecraft inner(convergent) Aaps of the exhaust nozzles are ex- as demonstrated at the prototype part level some years posed to higher temperatures(up to 1100C). Durabil go within the scope of the Hermes European space iry of the order of 1000 h has been demonstrated for shuttle project. Here, the maximum temperature 3D-composites with a self-healing multilayered matrix ranges from 800%C to 1600%C, during the ascent and on the basis of bench combustion tests. The next step is re-entry phases of a fight, the structures being submit- the combustion chamber or combustor, whose fabrica- ted to thermal shocks and cyclic mechanical loading tion with SiC-based composites is in progress. No re- under ablative or passive oxidizing atmospheres, with an sults of tests are presently available for military aerojet expected durability of a few tens of hours( Fig. 4a). Such engine combustors in the open literature, as far as we conditions are compatible with modern composites fab- know. However, the use of SiC-matrix combustor in ricated with carbon fibers(to reduce weight and achieve power plant gas turbine of cogeneration is well docu good mechanical properties at the highest temp mented, with similar (not to say more severe)service and engineered multilayered self-healing matrix. How- condition. 2.33 Combustors of large size comprising ever, some environmental barrier coating(EBC)might concentric cylindrical CMC liners have been fabricated undaccessary to limit the recession rate of the material by CVI or RMI with SiC(Hi-Nicalon)/BN/SiC(Si) under active oxidation/ablation regime(hT and veloc- composites. Durability of several 10,000 h has been es- ity combined with low P(O2). Such a material ap tablished under real service conditions, for composites proach will benefit from the high refractoriness of C/ with a BSAS-EBC.' To conclude, the use of SiC-ma- SiC composites(a 2500%C)relative to the low melting trix composites in the hot nonrotating parts of gas tur- point of aluminum(a 650%C) in the metallic option bines(aerojet engines and cogeneration gas turbines combined with a thermal insulation. Further, the CVI- appears to be promising. It is now a matter of engi process is well suited to the fabrication of large size neering, reliability, and cost(that of performant SiC structures(two meters or more), its feasibility already fibers still remaining relatively dissuasivefluid oxides and the in-depth diffusion of oxygen is slowed down or stopped.19 An even more efficient strat￾egy consists in replacing the homogeneous SiC-matrix itself by an engineered multilayered matrix based on a similar principle, referred to as a self-healing matrix and that could be infiltrated by P-CVI at the laboratory scale.27 Here, the matrix is deposited as the repetition of a given sequence S comprising thin layers X acting as mechanical fuses (X being C, C (B), BN, BN (Si), or any suitable fuse) and layers Y of species forming fluid ox￾ides, as mentioned above (Fig. 3b). Durability of the order of 1000 h under cyclic loading in air up to 11001C has been reported for SiC fiber composites fab￾ricated with such self-healing matrices.28 Finally, spe￾cific multilayered coatings containing oxide layers (such as mullite or/and baryum strontium aluminosilicate, BSAS) have been proposed and tested for SiC/SiC com￾posites exposed to wet oxidizing atmospheres to reduce the recession rate of the materials, e.g., in hot combus￾tion gas rich in water vapor.29 Representative Applications Space and Aeronautic Field SiC-matrix composites are potential material can￾didates for the fabrication of hot structures of spacecraft, as demonstrated at the prototype part level some years ago within the scope of the Hermes European space shuttle project.30 Here, the maximum temperature ranges from 8001C to 16001C, during the ascent and re-entry phases of a flight, the structures being submit￾ted to thermal shocks and cyclic mechanical loading under ablative or passive oxidizing atmospheres, with an expected durability of a few tens of hours (Fig. 4a). Such conditions are compatible with modern composites fab￾ricated with carbon fibers (to reduce weight and achieve good mechanical properties at the highest temperatures) and engineered multilayered self-healing matrix. How￾ever, some environmental barrier coating (EBC) might be necessary to limit the recession rate of the material under active oxidation/ablation regime (HT and veloc￾ity combined with low P(O2)). Such a material ap￾proach will benefit from the high refractoriness of C/ SiC composites (
25001C) relative to the low melting point of aluminum (
6501C) in the metallic option combined with a thermal insulation. Further, the CVI￾process is well suited to the fabrication of large size structures (two meters or more), its feasibility already established and joining techniques under development. Similar materials could also be used for the shielding of satellites against the impact of meteorites or foreign ob￾jects owing to the high toughness and hardness of these materials. Another promising field of application is that of the hot structures of aerojet engines and related gas turbines, which are presently made of heavy, low melting point nickel-based superalloys that require complex cooling systems. Replacing superalloys by SiC-based composites would permit to raise the gas temperature, suppress, or at least limit the cooling requirement, increase the effi- ciency of the engine, and reduce both the weight and the noise/pollution level (Fig. 4b). However, it will proba￾bly take some time (these materials are still young and their fabrication costly) and be limited in a first step to nonrotating parts, i.e., the combustors and the after￾burner parts such as the flaps of the exhaust nozzles. The main concern here is durability, that should be of the order of several thousands of hours. The outer (diver￾gent) flaps of exhaust nozzles experience a temperature that is relatively low (To7001C). Hence, they can be fabricated with carbon fiber-reinforced SiC-matrix, with a weight gain of 50%. Tested in flight as early as 1989, they are now in volume production Q3 (M88 Snecma en￾gines of the Rafale fighter, Paris, France).28,30,31 The inner (convergent) flaps of the exhaust nozzles are ex￾posed to higher temperatures (up to 11001C). Durabil￾ity of the order of 1000 h has been demonstrated for 3D-composites with a self-healing multilayered matrix on the basis of bench combustion tests. The next step is the combustion chamber or combustor, whose fabrica￾tion with SiC-based composites is in progress. No re￾sults of tests are presently available for military aerojet engine combustors in the open literature, as far as we know. However, the use of SiC-matrix combustor in power plant gas turbine of cogeneration is well docu￾mented, with similar (not to say more severe) service condition.32,33 Combustors of large size comprising concentric cylindrical CMC liners have been fabricated by CVI or RMI with SiC (Hi-Nicalon)/BN/SiC (Si) composites. Durability of several 10,000 h has been es￾tablished under real service conditions, for composites with a BSAS–EBC.29 To conclude, the use of SiC-ma￾trix composites in the hot nonrotating parts of gas tur￾bines (aerojet engines and cogeneration gas turbines) appears to be promising. It is now a matter of engi￾neering, reliability, and cost (that of performant SiC fibers still remaining relatively dissuasive). www.ceramics.org/ACT SiC-Matrix Composites: Application 81