• Journal of the Korean Society of Combustion
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• v.12 no.1
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• pp.11-20
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• 2007

The supersonic combustion experiments are carried out using T3 free-piston shock tunnel. Different shock tube fill pressures have various inflow conditions. $15^{\circ}$ inclined hydrogen fuel injection is located before the cavity. Oblique shock is generated from the cavity and reflects off the top and bottom wall. For non-reacting flow, fuel makes the shear layer thicker above the cavity therefore, the shock is generated just before the trailing edge. This research has self-ignition in the combustor. For reacting flow, as the equivalence ratio increases, flame starts to generate near the injector or occur in the recirculation zone before the injector. High fuel injection sustains the jet shape in the cross flow and air can mix with fuel along the shear layer. Therefore, two flame layers find above the cavity for high equivalence ratio.

• Fire Science and Engineering
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• v.27 no.6
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• pp.64-69
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• 2013

The Numerical simulation was performed on the flow field around the two-dimensional rectangular bluff body in order to simulate an engine nacelle fire and to complement the previous experimental results of the bluff body stabilized flames. Fire Dynamic Simulator (FDS) based on the Direct Numerical Simulation (DNS) was employed to clarify the characteristics of reacting flow around bluff body. The overall reaction was considered and the constant for reaction was determined from flame extinction limits of experimental results. The air used atmosphere and the fuel used methane. For both fuel ejection configurations against an oxidizer stream, the flame stability and flame mode were affected mainly by vortex structure near bluff body. In the coflow configuration, air velocity at the flame extinction limit are increased with fuel velocity, which is comparable to the experiment results. Comparing with the isothermal flow field, the reacting flow produces a weak and small recirculation zone, which is result in the reductions of density and momentum due to temperature increase by reaction in the wake zone.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.04a
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• pp.43-51
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• 1997

Steady and unsteady numerical simulations are conducted for the experiments performed to investigate the ram accelerator flow field by using the expansion tube facility in Stanford University. Navier-Stokes equations for chemically reacting flows are analyzed by fully implicit and time accurate numerical methods with Jachimowski's detailed chemistry model for hydrogen-air combustion involving 9 species and 19 reaction steps. Although the steady state assumption shows a good agreement with the experimental schlieren and OH PLIF images for the case of $2H_2$+$O_2$+$17N_2$, it fails in reproducing the combustion region behind the shock intersection point shown in the case of $2H_2$+$O_2$+$12N_2$, mixture. Therefore, an unsteady numerical simulation is conducted for this case and the result shows all the detailed flow stabilization process. The experimental result is revealed to be an instantaneous result during the flow stabilization process. The combustion behind the shock intersection point is the result of a normal detonation formed by the intersection of strong oblique shocks that exist at early stage of the stabilization process. At final stage, the combustion region behind the shock intersection point disappears and the steady state result is retained. The time required for stabilization of the reacting flow in the model ram accelerator is found to be very long in comparison with the experimental test time.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.27 no.1
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• pp.95-104
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• 2003

The objective of this work is to investigate the turbulent reacting flow in a three dimensional combustor with emphasis on thermal NO emission through a numerical simulation. Flow field is analyzed using the SIMPLE method which is known as stable as well as accurate in the combustion modeling, and the finite volume method is adopted in solving the radiative transfer equation. In this work, the thermal characteristics and NO emission in a three dimensional combustor by changing parameters such as equivalence ratio and inlet swirl angle have investigated. As the equivalence ratio increases, which means that more fuel is supplied due to a larger inlet fuel velocity, the flame temperature increases and the location of maximum temperature and thermal NO has moved towards downstream. In the mean while, the existence of inlet swirl velocity makes the fuel and combustion air more completely mixed and burnt in short distance. Therefore, the locations of the maximum reaction rate, temperature and thermal NO were shifted to forward direction compared with the case of no swirl.

• Journal of Mechanical Science and Technology
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• v.17 no.1
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• pp.139-145
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• 2003

Diode laser sensor is conducted to measure the gas temperature in the liquid-gas 2-phase counter flow flame. C$\_$10/H/ sub 22/ and city gas were used as liquid fuel and gas fuel, respectively. Two vibrational overtones of H$_2$O were selected and measurements were carried out in the spray flame region stabilized the above gaseous premixed flame. The path-averaged temperature measurement using diode laser absorption method succeeded in the liquid fuel combustion environment regardless of droplets of wide range diameter. The path-averaged temperature measured in the post flame of liquid-gas 2-phase counter flow flame showed qualitative reliable results. The successful demonstration of time series temperature measurement in the liquid-gas 2-phase counter flow flame gave us motivation of trying to establish the effective control system in practical combustion system. These results demonstrated the ability of real-time feedback from combustor inside using the non-intrusive measurement as well as the possibility of application to practical combustion system. Failure case due to influence of spray flame was also discussed.

• Journal of computational fluids engineering
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• v.12 no.1
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• pp.60-67
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• 2007

In this paper, the dispersion controlled dissipative (DCD) scheme is reviewed and then extended to simulate chemically reacting gas flows in multicomponent mixtures in the presence of strong shock waves. Furthermore, the properties of the reactive DCD (DCD-R) scheme are discussed, followed by several applications. The DCD scheme has been shown to have the following features: high accuracy and robustness for reacting gas flows in the presence of strong shock waves and contact discontinuities, and algorithmic simplicity.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.238-241
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• 2007

We present an innovative method of multi-physics application involving energetic materials. Energetic materials are related to reacting flows in extreme environments such as fires and explosions. They typically involve high pressure, hish temperature, strong non-linear shock waves, and high strain rate deformation of metals. We use an Eulerian methodology to address these problems. Our approach is naturally free from large deformation of materials that makes it suitable for high strain-rate multi-material interaction problems. Furthermore we eliminate the possible interface smearing by using the level sets. We have devised a new level set based tracking framework that can elegantly handle large gradients typically found in reacting gases and metals. We show several work-in-progress applications of our algorithm including the Taylor impact test, explosive venting and additional confined explosion problems of modem interest.

• Journal of Advanced Marine Engineering and Technology
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• v.18 no.1
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• pp.81-91
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• 1994

Combustion noise involved with chemical heat release and turbulent process in turbopropulsion systems, gasturbine, industrial furnaces and internal engines is indeed noisy. The experimental study reported in this paper is made to identify a dominant combustion noise in jet flames. Gaseous propane and kerosene fuel have been used with air as the oxidizer in a different jet combustion systems. Combustion and aerodynamic noise are studied through far field sound pressure measurements in an anechoic chamber. And also mean temperature and velocities and turbulent intensities of both isothermal and reacting flow fields were measured. It is shown that axial mean velocity of reacting flow fields is higher about 1 to 3m/sec than that of cold flow in a gaseous combustor. As the gaseous fuel flow rate increases, the acoustic power increases. But the sound pressure level for the spray flame decreases with increasing equivalence ratio. The influence of temperature in the combustion fields due to chemical heat release has been observed to be a dominant noise source in the spray flame. The spectra of combustion noise in gaseous propane and kerosene jet flame show a predominantly low frequency and a broadband nature as compared with the noise characteristics in an isothermal air jet.

• 한국연소학회:학술대회논문집
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• 2006.10a
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• pp.209-219
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• 2006

The supersonic combustion experiments are carried out using T3 free-piston shock tunnel. Different shock tube fill pressures have various inflow conditions. $15^{\circ}$ inclined hydrogen fuel injection is located before the cavity. Oblique shock is generated at the trailing edge of the cavity and reflects off the top and bottom wall. For non-reacting flow, static pressures in low equivalence ratio are similar to those in no fuel injection. As equivalence ratio is increased, static pressures are increased in the duct. In the similar equivalence ratio, static pressures are increased when total enthalpy is decreased. For reacting flow, the flame is occurred near the cavity. The combustion is weak locally in the middle of the duct. The up and down pressure distribution in the duct means that the supersonic combustion is generated.

• Journal of computational fluids engineering
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• v.6 no.2
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• pp.32-39
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• 2001

Hypersonic analysis on the KSR-Ⅲ payload configuration has been performed using an axisymmetric Navier-Stokes code. A numerical code based on the Harten and Yee's upwind TVD scheme with simplified curve fits in the chemically reacting equilibrium air was developed. The carbuncle phenomenon on detached shock in front of the payload is controlled by using pressure gradients to tune the dissipation. Chemically reacting equilibrium computations for the reentry flight conditions of Mach No. 10.2, 8, 4.9 are presented and compared with the results of calorically perfect gas.