G.N. Morscher et al/ Composites Science and Technology 68 (2008)3305-3313 greater than 1%)are achieved. However, times to failure in air are Some of the elevated temperature tests were performed in usually less than 100 h for stresses that are in excess of the matrix other studies[9-10. All the tests were performed at 1204C in cracking stress (e.g, x100 MPa for the Hi-Nicalon m CVI SiC com- lab air at Southern research Institute, Birmingham AL or Cincinnati posite tested in Ref. 4]at 1300C Test Labs, Cincinnati, OH. Tensile creep tests were performed under For SiC/SiC composites reinforced with high modulus poly constant load using either universal testing or lever-arm machines talline SiC fibers(Hi-Nicalon S)and a slurry-cast melt-infiltrated The cyclic fatigue tests were all performed in hydraulic testing ma- (MI) SiC matrix composite system a somewhat different damage chines; while dwell fatigue testing was done in modified lever-arm development was observed at 1315C[7. For the stress ranges tensile machines to control loading rates. Several fatigue loading tested(up to 172 MPa), only minor microcracking in the 90 mini- rates were used. Dwell fatigue(DF)consisted of a 2 h hold followed composites was observed. In some cases, these cracks were by a 1 min unload-reload back to a two hour hold High cycle fati- observed to extend to the surface(138 MPa)where 90 minicom- gue(HCF)was performed at 1 or 30 Hz. All of the fatigue tests were osites were adjacent to the surface and into some 0 minicompos- performed for an R ratio of 0.05. Displacement for all tests was tes at higher stresses(172 MPa)resulting in local fiber failure but monitored using contact extensometers on the edge of the speci- not significant through-the-thickness matrix crack formation. men with a gage length of 25 mm. Creep strain was determined 00 h followed by determination of retained stress-strain behavior, perature along the gage section he furnace assured uniform tem hich was usually very high since minimal damage occurred in the For some of the specimens which were not taken to failure, a opposites. room temperature unload-reload hysteresis tensile test was per It is the goal of this study to further understand the develop- formed to failure to determine residual properties. aE sensors were ment of damage in a similar polycrystalline Sic fiber-reinforced placed above and below the extensometers(50-60 mm apart ) and slurry-cast MI SiC matrix system, the Sylramic-iBN MI SiC sys- AE was monitored during the room temperature test using a Digi em developed at NASA Glenn Research Center and referred to tal Wave Fracture Wave detector (Englewood Co)as described as N24A[ 8]. This material has undergone some of the most elsewhere [7, 11. After the test, the events were sorted as to loca exhaustive testing to date of any high-performance SiC/Sic com- tion along the specimen length and only those events which posite system under an US Air Force sponsored program, includ- occurred in the 25 mm gage section were used for AE analysis. tensile creep, 2 hour dwell fatigue, 1 Hz cyclic fatigue, and Analysis of the crept/fatigued specimens consisted of both opti 30 Hz cyclic fatigue at 1204C in air for stresses ranging from cal and scanning electron microscopy. Fig. 1 shows a typical failed 110 to 220 MPa for times up to 2000h [ 9-10. Specimens from specimen. One part of the fracture surface was used for observation this wide range of tests were acquired for this study in order of the fracture surface in a field emission scanning electron micro- to determine the damage accumulation for the wide range of scope(FESEM-Hitachi 4700, Tokyo Japan). The other half of the stress-time conditions which will be described here. Some of specimen was cut(between 10 and 15 mm long) and polished the specimens had been interrupted at predetermined times. along the edge (approximately 1 mm from the exposed edge For most of those specimens, a room temperature unload-reload and or face of the specimen in order to observe and quantify the test to failure was performed with acoustic emission(AE)mon- number of matrix cracks along the length Matrix cracks caused itoring in the same way as Ref. [7 in order to determine the by time-dependent deformation had significant crack-openings residual properties of the composite and were easy to distinguish from matrix cracks formed during the room temperature retained strength tests, which were often not discernable as-polished due to crack closure from the high 2. Experimental compressive stresses in the matrix For the edge-polished speci- mens, matrix cracks were counted along the length for both The composite system evaluated is described in more detail elsewhere [8 It consists of eight plies of 2D woven five-harness satin Sylramic-iBN fabric, a CVI BN fiber interphase coating, a CVI Sic coating of initial matrix to rigidize the preform and protect the fibers, slurry-infiltrated SiC particulates, and molten Si infil- trated to fill in the remaining open porosity. The initial five-harness fabric consisted of 7.9 tow ends per cm of Sylramic SiC fiber bal anced in the warp and weft 0/90 orthogonal directions. The fiber ows consisted of 800 fibers approximately 10 um in diameter. The fibers were originally produced by Dow Corning, Midland, Michigan, but are now produced by ATK Col Ceramics in San Diego CA The precursor Sylramic fabric plies were subjected to a NASA- proprietary treatment in order to improve fiber creep-resistance and also produce a thin(150 nm) in situ grown BN layer on the sEM|丙 urface of each fiber prior to composite fabrication. The composites were fabricated as 153 x 230 mm panels by ply lay-up at GE Cera- nic Composite Products, LLC in Newark, DE All plies were aligned /90 in plane but were randomly stacked through-the-thickness with a degree of fiber nesting. The fiber volume fraction of compos- ites was approximately 36-38% as measured by weight. The tensile specimens were machined from the as-fabricated panels with a length of 155 mm, a grip width of 12 mm, and a dogboned section in the middle with a length of 40 mm and width of 8.2 mm. One of the orthogonal fiber directions was aligned within +/-3 of the specimen length or tensile direction. Typical specimen thickness was 2 mm. Fig. 1. Typical failed specimen after creep or fatigue.greater than 1%) are achieved. However, times to failure in air are usually less than 100 h for stresses that are in excess of the matrix cracking stress (e.g., 100 MPa for the Hi-NicalonTM CVI SiC com￾posite tested in Ref. [4] at 1300 C). For SiC/SiC composites reinforced with high modulus polycrys￾talline SiC fibers (Hi-Nicalon S) and a slurry-cast melt-infiltrated (MI) SiC matrix composite system a somewhat different damage development was observed at 1315 C [7]. For the stress ranges tested (up to 172 MPa), only minor microcracking in the 90 mini￾composites was observed. In some cases, these cracks were observed to extend to the surface (138 MPa) where 90 minicom￾posites were adjacent to the surface and into some 0 minicompos￾ites at higher stresses (172 MPa) resulting in local fiber failure but not significant through-the-thickness matrix crack formation. However, for this study, creep times were usually limited to 100 h followed by determination of retained stress-strain behavior, which was usually very high since minimal damage occurred in the composites. It is the goal of this study to further understand the develop￾ment of damage in a similar polycrystalline SiC fiber-reinforced slurry-cast MI SiC matrix system, the Sylramic-iBN MI SiC sys￾tem developed at NASA Glenn Research Center and referred to as N24A [8]. This material has undergone some of the most exhaustive testing to date of any high-performance SiC/SiC com￾posite system under an US Air Force sponsored program, includ￾ing tensile creep, 2 hour dwell fatigue, 1 Hz cyclic fatigue, and 30 Hz cyclic fatigue at 1204 C in air for stresses ranging from 110 to 220 MPa for times up to 2000 h [9–10]. Specimens from this wide range of tests were acquired for this study in order to determine the damage accumulation for the wide range of stress-time conditions which will be described here. Some of the specimens had been interrupted at predetermined times. For most of those specimens, a room temperature unload-reload test to failure was performed with acoustic emission (AE) mon￾itoring in the same way as Ref. [7] in order to determine the residual properties of the composite. 2. Experimental The composite system evaluated is described in more detail elsewhere [8]. It consists of eight plies of 2D woven five-harness satin Sylramic-iBN fabric, a CVI BN fiber interphase coating, a CVI SiC coating of initial matrix to rigidize the preform and protect the fibers, slurry-infiltrated SiC particulates, and molten Si infil￾trated to fill in the remaining open porosity. The initial five-harness fabric consisted of 7.9 tow ends per cm of SylramicTM SiC fiber bal￾anced in the warp and weft 0/90 orthogonal directions. The fiber tows consisted of 800 fibers approximately 10 lm in diameter. The fibers were originally produced by Dow Corning, Midland, Michigan, but are now produced by ATK COI Ceramics in San Diego, CA. The precursor Sylramic fabric plies were subjected to a NASA￾proprietary treatment in order to improve fiber creep-resistance and also produce a thin (150 nm) in situ grown BN layer on the surface of each fiber prior to composite fabrication. The composites were fabricated as 153 230 mm panels by ply lay-up at GE Cera￾mic Composite Products, LLC in Newark, DE. All plies were aligned 0/90 in plane, but were randomly stacked through-the-thickness with a degree of fiber nesting. The fiber volume fraction of compos￾ites was approximately 36–38% as measured by weight. The tensile specimens were machined from the as-fabricated panels with a length of 155 mm, a grip width of 12 mm, and a dogboned section in the middle with a length of 40 mm and a width of 8.2 mm. One of the orthogonal fiber directions was aligned within +/3 of the specimen length or tensile direction. Typical specimen thickness was 2 mm. Some of the elevated temperature tests were performed in other studies [9–10]. All the tests were performed at 1204 C in lab air at Southern Research Institute, Birmingham AL or Cincinnati Test Labs, Cincinnati, OH. Tensile creep tests were performed under constant load using either universal testing or lever-arm machines. The cyclic fatigue tests were all performed in hydraulic testing ma￾chines; while dwell fatigue testing was done in modified lever-arm tensile machines to control loading rates. Several fatigue loading rates were used. Dwell fatigue (DF) consisted of a 2 h hold followed by a 1 min unload-reload back to a two hour hold. High cycle fati￾gue (HCF) was performed at 1 or 30 Hz. All of the fatigue tests were performed for an R ratio of 0.05. Displacement for all tests was monitored using contact extensometers on the edge of the speci￾men with a gage length of 25 mm. Creep strain was determined after reaching full applied stress. The furnace assured uniform tem￾perature along the gage section. For some of the specimens which were not taken to failure, a room temperature unload-reload hysteresis tensile test was per￾formed to failure to determine residual properties. AE sensors were placed above and below the extensometers (50–60 mm apart) and AE was monitored during the room temperature test using a Digi￾tal Wave Fracture Wave detector (Englewood CO) as described elsewhere [7,11]. After the test, the events were sorted as to loca￾tion along the specimen length and only those events which occurred in the 25 mm gage section were used for AE analysis. Analysis of the crept/fatigued specimens consisted of both opti￾cal and scanning electron microscopy. Fig. 1 shows a typical failed specimen. One part of the fracture surface was used for observation of the fracture surface in a field emission scanning electron micro￾scope (FESEM – Hitachi 4700, Tokyo Japan). The other half of the specimen was cut (between 10 and 15 mm long) and polished along the edge (approximately 1 mm from the exposed edge) and/or face of the specimen in order to observe and quantify the number of matrix cracks along the length. Matrix cracks caused by time-dependent deformation had significant crack-openings and were easy to distinguish from matrix cracks formed during the room temperature retained strength tests, which were often not discernable as-polished due to crack closure from the high compressive stresses in the matrix. For the edge-polished speci￾mens, matrix cracks were counted along the length for both Fig. 1. Typical failed specimen after creep or fatigue. 3306 G.N. Morscher et al. / Composites Science and Technology 68 (2008) 3305–3313