• The Journal of Korean Academy of Prosthodontics
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• v.53 no.4
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• pp.295-300
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• 2015

Purpose: The aim of this study was to evaluate the shear bond strength between Ni-Cr alloy and composite resin using universal adhesive systems coMPared to conventional method using metal primers. Materials and methods: For this study, a total of 120 cast commercial Ni-Cr alloy (Vera Bond 2V) disks were embedded in acrylic resin, and their surfaces were smoothed with silicon carbide papers and airborne-particle abrasion. Specimens of each metal were divided into 6 groups based on the combination of metal primers (Metal primer II, Alloy primer, Metal & Zirconia primer, MKZ primer) and universal adhesive systems (Single Bond Universal, All Bond Universal). All specimens were stored in distilled water at $37^{\circ}C$ for 24 hours. Shear bond strength testing was performed with a universal testing machine at a cross head speed of 1 m/min. Data (MPa) were analyzed using one-way ANOVA and the post hoc Tukey's multiple comparison test (${\alpha}$=.05). Results: There were significant differences between Single Bond Universal, All Bond Universal, Metal Primer II and Alloy Primer, MKZ Primer, Metal & Zirconia Primer (P<.001). Conclusion: Universal Adhesive system groups indicated high shear bond strength value bonded to Ni-Cr alloy than that of conventional system groups using primers except Metal Primer II. Within the limitations of this study, improvement of universal adhesive systems which can be applied to all types of restorations is recommended especially non-precious metal alloy. More research is needed to evaluate the effect of silane inclusion or exclusion in universal adhesive systems.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.91-93
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• 2003

A comprehensive numerical study is carried out to investigate for the understanding of the flow evolution and flame development in a supersonic combustor with normal injection of ncumally injecting hydrogen in airsupersonic flows. The formulation treats the complete conservation equations of mass, momentum, energy, and species concentration for a multi-component chemically reacting system. For the numerical simulation of supersonic combustion, multi-species Navier-Stokes equations and detailed chemistry of H2-Air is considered. It also accommodates a finite-rate chemical kinetics mechanism of hydrogen-air combustion GRI-Mech. 2.11[1], which consists of nine species and twenty-five reaction steps. Turbulence closure is achieved by means of a k-two-equation model (2). The governing equations are spatially discretized using a finite-volume approach, and temporally integrated by means of a second-order accurate implicit scheme (3-5).The supersonic combustor consists of a flat channel of 10 cm height and a fuel-injection slit of 0.1 cm width located at 10 cm downstream of the inlet. A cavity of 5 cm height and 20 cm width is installed at 15 cm downstream of the injection slit. A total of 936160 grids are used for the main-combustor flow passage, and 159161 grids for the cavity. The grids are clustered in the flow direction near the fuel injector and cavity, as well as in the vertical direction near the bottom wall. The no-slip and adiabatic conditions are assumed throughout the entire wall boundary. As a specific example, the inflow Mach number is assumed to be 3, and the temperature and pressure are 600 K and 0.1 MPa, respectively. Gaseous hydrogen at a temperature of 151.5 K is injected normal to the wall from a choked injector.A series of calculations were carried out by varying the fuel injection pressure from 0.5 to 1.5MPa. This amounts to changing the fuel mass flow rate or the overall equivalence ratio for different operating regimes. Figure 1 shows the instantaneous temperature fields in the supersonic combustor at four different conditions. The dark blue region represents the hot burned gases. At the fuel injection pressure of 0.5 MPa, the flame is stably anchored, but the flow field exhibits a high-amplitude oscillation. At the fuel injection pressure of 1.0 MPa, the Mach reflection occurs ahead of the injector. The interaction between the incoming air and the injection flow becomes much more complex, and the fuel/air mixing is strongly enhanced. The Mach reflection oscillates and results in a strong fluctuation in the combustor wall pressure. At the fuel injection pressure of 1.5MPa, the flow inside the combustor becomes nearly choked and the Mach reflection is displaced forward. The leading shock wave moves slowly toward the inlet, and eventually causes the combustor-upstart due to the thermal choking. The cavity appears to play a secondary role in driving the flow unsteadiness, in spite of its influence on the fuel/air mixing and flame evolution. Further investigation is necessary on this issue. The present study features detailed resolution of the flow and flame dynamics in the combustor, which was not typically available in most of the previous works. In particular, the oscillatory flow characteristics are captured at a scale sufficient to identify the underlying physical mechanisms. Much of the flow unsteadiness is not related to the cavity, but rather to the intrinsic unsteadiness in the flowfield, as also shown experimentally by Ben-Yakar et al. [6], The interactions between the unsteady flow and flame evolution may cause a large excursion of flow oscillation. The work appears to be the first of its kind in the numerical study of combustion oscillations in a supersonic combustor, although a similar phenomenon was previously reported experimentally. A more comprehensive discussion will be given in the final paper presented at the colloquium.