Ken Goto

第3回観測ロケットシンポジウム（2021年3月24-25日. オンライン開催) 3rd Sounding Rocket Symposium (March 24-25, 2021. Online Meeting) 著者人数: 14名PDF再処理の為、2023年3月17日に差替 資料番号: SA6000162017 レポート番号: Ⅴ-3

第35回宇宙構造・材料シンポジウム（2019年12月2日. 宇宙航空研究開発機構宇宙科学研究所 (JAXA)(ISAS)), 相模原市, 神奈川県資料番号: SA6000146015レポート番号: A-14

ZrB_2-SiC-ZrC(ZSZ) is known as one of ultra-high temperature ceramics (UHTCs) composite materials. In this study, four different composition of the ZSZ and ZrB_2-SiC were fabricated by spark plasma sintering. The specimens were oxidized at 1700℃ for 10 min in two different oxygen partial pressure. After the oxidation tests, SiO_2 and ZrO_2 were formed on the surface of both ZSZs and ZrB_2-SiC. In addition, unoxidized region still remained under these oxides. In case of the ZSZ, thickness of oxidized layer increased with ZrC volume content. After oxidation test under low oxygen partial pressure, porous layer was observed in the ZrB_2-SiC, but no such layer was observed in the ZSZs. Since the porous layer is brittle and low strength, the ZSZ has an advantage in the mechanical properties after oxidation. In order to validate the advantage of the ZSZ, the oxidation behavior was thermodynamically analyzed using volatility diagrams.

In the present study, long term durability of the candidate material for the next radio astronomy satellite under space environment were investigated. The materials were high modulus PAN based carbon fiber reinforced polycyanate ester and epoxy. In order to simulate the space environment, thermal cycle tests between -197 and 120 ℃ were conducted. Surface observations and bending modulus measurements were also conducted. Transverse cracks in 90 ° plies were observed at 1 cycle and the number of crack increased with cycles. Bending modulus, however remain constant during thermal cycles. Microscopic damage in carbon fiber reinforced polycyanate ester was suppressed at the late stage of the thermal cycle. This result indicated the long term durability for carbon fiber reinforced polycyanate ester was superior.

ZrB_2-SiC-ZrC (ZSZ), which is known as one of ultra high temperature ceramics (UHTCs) composite material, with four different compositions containing 16 vol.%SiC and ZrB_2-16 vol.%SiC were sintered by spark plasma sintering. These sintered compacts are oxidized at 1700 ℃ by IR image furnace in airflow. The result showed that all samples were covered with SiO_2 and ZrO_2. Moreover, oxide scale of ZSZs containing higher ZrC than ZrB_2 has been damaged. Observation of the cross-section of all samples revealed that oxide area was consist of five different layers: SiO_2 layer, SiO_2 with ZrO_2 layer, SiC-depleted layer and unoxidized area. Furthermore, the thickness of oxide layer of ZSZs containing more ZrB_2 than ZrC was thinner. Thus, it was expected that these samples have good resistance to oxidation.

Solar power sail that is being developed by JAXA as a planetary exploration spacecraft utilizes a combination of electric and photon propulsion system. The solar power sail requires a super lightweight flexible power generation system that uses thin film solar cells. We are developing a mounting method of the thin film solar cells on the membrane of the sail. Major issues of our study are shape control and connection between thin film solar array and thin film electrical wiring harnesses. We describe experimental results of shape control by oxide sputtering on the polyimide film and thin film solar cells and evaluation results of trial production of the thin film solar array connected with the thin film flexible wiring harnesses by heat sealing.

Ultra thin fiber reinforced plastic plates, with around l0μm thickness, were fabricated using the uni-directionally aligned carbon nanotube (CNT) sheet. The CNT sheets are drawn wound around the bobbin from the CNT array developed on the silica glass base plate. The CNT used in this study was multi-walled CNT with diameter of around 40 nm. The CNT/epoxy resin prepreg was created by heating the CNT sheet put on the epoxy resin sheet at the temperature of 100℃ for 3 minutes using a hotpress. The CNT/epoxy resin composite material is laminated on the symmetrical cross ply lamination ([90/0]s). To minimize the content of epoxy resin to have thinner thickness, the CNT sheet with out resin were used to 0° layers and the prepreg to 90° layers. The total thickness of the composite became around 40 μm with high volume fraction of ≈ 30 to 50 %. However, the modulus and strength of the composites in tensile test became low. The low modulus attributes from the ply thickness un-uniformity where 0° ply thickness became about 1.5 times larger than 90° ply. Because of this thickness un-uniformity, CNT volume fraction in 0° plies became small to make tensile modulus low To improve mechanical properties of thin CNT composite film, process optimization to have thinner and uniform ply thickness were required as a future work.

ZrB_2-SiC-ZrC(ZSZ) with four different compositions containing 16 vol% SiC were sintered by hot pressing and spark plasma sintering. In this study, ZSZs were oxidized in the ultra-high temperature range by two different tests. Firstly, ZSZs were oxidized at 1973 K in an airflow by an IR image furnace for several minutes. The result showed that ZSZs have oxidation resistance in this temperature range due to formation of oxide scale which acts as a barrier against oxygen diffusion into the unoxidized area. Moreover, ZSZ containing more ZrC formed ZrO_2 rich scale, whereas ZSZ containing more ZrB_2 formed SiO_2 rich scale. These results suggested that the highest oxidation resistance would berealized if the silica scale is formed with ZrO_2 rigid skelton. In the next step, ZSZ, ZrB_2-SiC and monolithic SiC were oxidized continuously up to 2073 K in air by a ZrO_2 furnace to compare their oxidation resistance. The results revealed that oxide scale disappeared from monolithic SiC surface, whereas ZSZ and ZrB_2-SiC maintained their oxide scaleduring the oxidation test. Comparing ZSZ with ZrB_2-SiC, ZSZ formed silica scale with ZrO_2 skelton, whereas no ZrO_2 skelton was formed in the silica scale of ZrB_2-SiC. The IR image furnace test implied that it is important to have not only high oxidation resistant SiO_2 layer but also coexistence rigid oxide scale. Therefore, it is suggested that ZSZ is more attractive candidate for the advanced aerospace heat-resistant material than SiC and ZrB_2-SiC.

For a high accuracy antenna in next radio astronomy satellite, a candidate material is Carbon Fiber Reinforced Plastics (CFRP). Because longitudinal thermal expansion coefficient of unidirectional CFRP is negative, we can design a laminate without thermal expansion. This enables high structural accuracy under large temperature fluctuation like space. On the other hand when the laminate is subjected to thermal impact, thermal stress occurs and causes microscopic damages. In this study, we characterize damage progress in CFRP laminates under cyclic thermal impact. Specimens were subjected to thermal cycle from -197℃to120℃. Thermal cycle is periodically stopped for surface observation and flexural loading. Transverse cracks accumulated with thermal cycle and saturated at about 80 cycles. Flexural modulus decreased at l cycle and remained the same.

ZrB_2-SiC-ZrC (ZSZ) ceramics were oxidized by two different methods to evaluate their oxidation mechanism and oxidation resistance at ultra high temperature region. Firstly, ZSZs were oxidized at 1700℃ by IR image furnace. Experimental results revealed that the glass formed on specimens with ZrB_2 rich composition was flown during oxidation and removed from the surface by the airflow. These results lead to the depletion of oxidation resistance because it is attribute to formation of the glass,. In addition, ZSZs were also oxidized continuously up to 1700℃ using a zirconia ultra high temperature furnace. The result showed that the ZSZs formed oxide scale, and the specimen after the experiment still had unoxidized area. These results imply that ZSZs have resistance of the active oxidation of the SiC. The results of two oxidation tests predict that ZSZs have optimal composition, and that the oxidation mechanism at ultra high temperature region should be taken into account to determine the optimal composition.

Deformation behaviors of TBC system under four major modes of thermo-mechanical fatigue test conditions have been studied, i.e., in-phase and out-of-phase conditions under stress or strain control modes. Stress-strain behavior of TBC system under each thermo-mechanical fatigue condition was compared qualitatively considering elastic, creep and thermal strains. Micro-scale deformation behaviors of TBC system under thermo-mechanical fatigue conditions were discussed using calculated macro-scale behaviors.

ZrB_2-SiC-ZrC (ZSZ) ceramics with four different compositions containing 16 vol% SiC were sintered by hot press and spark plasma sintering. The relative density of all the ZSZ ceramics made by both methods reached around 95%. These ZSZs oxidized continuously up to 1500℃ using thermo gravimetric analysis. The ZSZ containing the highest ZrB_2 indicated the lowest weight gain, and that containing the highest ZrC indicated the highest weight gain. After the oxidization, the microstructure of the samples' surface and cross section were studied by X-ray diffraction analysis, optical microscope and field emission scanning electron microscope along with X-ray dispersive spectroscopy. Results of these analyses for the samples revealed that ZSZs have oxidation resistance due to formation of the oxide scale, which act as the barrier for oxygen diffusion into the unoxidized substrate. In the next study, ZSZ ceramics was heated continuously up to 1700℃ in air using zirconia ultra high temperature furnace. Oxidation test was earned out at 1700℃ by IR image furnace. The surface of the oxidized specimen had a number of bubbles and appeared to be covered with zirconia and glassy compounds. These results suggest that the active oxidation of SiC was occurred at the interface between the oxide scale and the substrate. In addition, because the spallation between the oxide scale and the substrate was not occurred, it was implied that ZSZs showed the resistance to the active oxidation of SiC.

A nozzle and a combustion chamber of a small liquid fuel rocket engine used as an upper stage engine of solid propellant rocket and an apogee engine of a satellite requires to operate without cooling by cryogenic fuels. From this reason, materials that can be used at high temperature in oxidizing environment are highly required for such thrusters. To meet with this demand, trial manufacture of a thruster made of SiC fiber-reinforced SiC matrix composite are carried out in this study. A SiC/SiC composite thruster was successfully manufactured by CVI and PIP combination process. The thruster passed the proof pressurize test up to 5.0 MPa (2.5 times higher than operation pressure, 2.0 MPa). However, First trial ended in an explosion just after the ignition of the engine. By changing engine start sequence (provide fuel and oxidizer just before the ignition to avoid penetration of the liquid into the pores), 30s continuous oneration of the engine succeeded.

The formation technology into complicated shapes is often required for the applications of C/Cs to hot structures. A newly devised hybrid bonding method was proposed as a bonding technique for C/Cs, which is easily applicable and results in high strength. This bonding is composed of carbon bonding infiltrated with Si melt in the cracks in bonding layer. Thus the hybrid bonding is formed without pressurization. Optimum process conditions of the hybrid bonding were first explored. Then using the optimized procedure, the strength of the hybrid bonding using 2D-C/Cs as substrates was evaluated at temperatures up to 2000K. The strength of the hybrid bonding increased with rise in test temperature. This bonding strength enhancement was shown to be caused by the presence of SiC within bonding layer (crack arrester) and the release of thermal mismatch stress between the bonding layer and substrate C/Cs.

Spin burst tests of carbon-carbon (C/C) rings were carried out to confirm fracture criterion under high speed rotation as a function of the inner/outer diameter ratio R. The C/C rings were of a quasi-isotropic lamination type and were confirmed to have isotropic strength by static tensile tests. The burst rotation speeds of the C/C rings were found to be much lower than those expected from stress distributions. Detailed inspection of stress distributions in the C/C rings revealed that rotational unbalance inherent in C/Cs caused high average hoop stress and stress concentration at around inner radius. These effects were shown to lower fracture rotation speeds.

Tensile strength and creep behavior of a 2D laminate carbon-carbon composite (C/C) were examined from room temperature to 2773 K in an inert atmosphere. The tensile strength of the C/C was monotonically enhanced with increasing test temperatures. In particular, significant improvement was observed at temperatures higher than 1773 K. In this temperature range, nonlinear stress-strain curves were observed at low deformation rates, but with increasing test speed, the stress-strain curves became linear until total fracture. The source of the apparent nonlinearity was thus concluded to be creep deformation, which appeared from 1773 K. Two ruling mechanisms for the strength enhancement of the C/C at elevated temperatures were identified. The first source was de-gassing of absorbed water, which had a dominant influence on the strength enhancement up to 1773 K. The second was creep deformation. This phenomenon was notable at temperatures higher than 1773 K, and produced much larger enhancement than the de-gassing.

Feasibility studies were carried out aiming at the application of carbon/carbon (C/C) composites to a turbine disk, heat exchangers, and a plug nozzle for an ATREX engine intended for use in a future TSTO space vehicle. In these applications, the maximum temperature was estimated to be about 1500℃. In order to withstand this high temperature, attempts were made to utilize three-dimensionally reinforced C/C composites. The most serious problem encountered in the application of C/Cs to the turbine disk was the loss of fragments of the composite located near the outer periphery due to strong centrifugal force, which resulted in severe vibration due to rotational imbalance. The heat exchangers and plug nozzle have complex shapes in order to realize a large heat exchanging area. Joined structures were explored for these components. The principal effort in these applications has been placed on finding structures requiring low joining strength and developing materials with low gas leakage.

Cold spin tests of three dimensional fiber-reinforced carbon-carbon composites were conducted for the turbine disk of the air-turbo ram jet engine (ATREX). The fly-out objects were detected at the rotation speed of 12000 r.p.m. and the fly-out behavior was caused by the delamination of the fiber bundles from the surface of the disk. The energy release rate, G, was estimated about 30 J/m2 for this 3D-C/C disk and fine fiber weave structure and large bonding strength between fiber bundle are shown to be effective to increase rotation speed. Gaseous Si conversion and Si infiltration method were examined to increase rotation speed of 3D-C/C disk.

The notch insensitivity of carbon-carbon composites (C/Cs) has been believed to result primarily from shear damage near sources of stress concentration. To evaluate this hypothesis, notch sensitivity has been examined for C/Cs with crossply laminates (CP-C/Cs) and quasi-isotropic (QI-C/Cs) laminates, The main difference in both laminates involves their shear behavior: the QI-C/Cs have an almost-linear stress-strain curve and high strength, whereas the CP-C/Cs exhibit strong nonlinearity and low strength. Thus, the effect of shear damage can be extracted by comparison of both materials. Experimental results from the present study have shown that the fracture behaviors of both C/Cs are quite similar. Finite-element analyses also have revealed that the stress redistribution caused by shear nonlinear deformation is too small to explain its toughening behavior, even in the CP-C/C, From these results, it is concluded that the notch-insensitive behavior of the C/Cs cannot be explained by the already-proposed shear-damage mechanism. To this end, a discussion has been conducted on a new possible toughening mechanism that is capable of generating the R-curve and notch insensitivity.

Compressive behavior of 2D and 3D Carbon/Carbon (C/C) composite was investigated. 2DC/C showed relatively linear deformation up to the total fracture. However, 3DC/C showed large nonlinear deformation accompanying the load drop and recovery and no clear fracture was observed. In 2D specimen, the progressive microcracking was not observed during the testing, however, interlaminar delamination was observed. In 3D specimen, the microcracks didn't appear until the first stress peak of the δ-U curve, and after the peak, the delamination between X-Y bundle and shear failure of the Y bundle near the matrix pocket was observed. From the experimental observation, the initiation of the progressive fracture of 3DC/C was supposed as the interlamina delamination, which ruled the fracture of 2DC/C. However, the existence of Z directional fiber prevented the specimen from catastrophic failure. The fracture strength of both the C/Cs was discussed using micro-kinking model by Rosen and shear fracture model by Ewins & Ham.

Interface debonding and its effect on crack stability in a single-edge-notched PMMA/glass bi-material have been studied. Direct observation of crack growth behavior is carried out and crack tip stress intensity factor during crack growth process is obtained by a laser Caustics method. Interface stress distribution at the onset of interface debonding is obtained by FEA with experimentally obtained stress/strain conditions. The interface debonding ahead of a growing crack tip occurs before the crack reaches the interface. A transition from slow to rapid crack growth is predicted by the event of an interface debonding ahead of a growing crack tip. This debonding is caused by an interface tensile stress. The maximum interface tensile stress obtained by FEA with experimental boundary condition, sigma(d), and the distance between the crack tip and interface at the onset of interface debonding, x(0), follows the relation, sigma(d) = A(d)/root x(0). This relation provides the condition for the interface tensile debonding. The condition suggests that the interface tensile debonding has a probabilistic nature.