• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.4
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• pp.342-348
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• 2017

In this paper, film cooling characteristics at supersonic free stream conditions were examined experimentally by applying an IR-thermography. Film cooling experiments were carried out in a free-jet facility at Mach number of 3.0 and with unit Reynolds number of $42.53{\times}10^6$ and $69.35{\times}10^6$ using wedge shaped film cooling model which has a converging film cooling nozzle. Film cooling efficiency was calculated by measuring the surface temperature of PEEK(Polyether Ether Ketone) and the effects of angle of attack and blowing ratios on the film cooling efficiency were examined. The measured wall temperature was significantly reduced by the film cooling flow compared with the results without the film cooling flow. The usefulness of film cooling was also confirmed by the surface heat flux calculated using the surface temperature history of PEEK. As the blowing ratio increases the protected area of PEEK was also expanded along the direction of free stream and film cooling flow.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.629-634
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• 2011

A prestudy on expendable turbine engine for high-speed vehicle was conducted. The two possible mission profiles were established to decide the engine requirements and Design Point, and Design Point analysis was performed with the values of design parameter which were obtained from similar class engines and technical references. The results showed that Specific Net Thrust is 2599.4 ft/s and Specific Fuel Consumption is 1.483 lb/($lb^*h$) at the flight condition of Sea Level, Mach 1.2. It was also found through the performance analysis on the two possible mission profiles that major design parameters for determining Net Thrust were Turbine Inlet Temperature for low supersonic flight speed and Compressor Exit Temperature for high supersonic flight speed. In addition, simple turbojet engine with axial compressor, straight annular combustor, axial turbine and fixed throat area converge-diverge exhaust nozzle was proposed as the configuration of simple low cost light engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.573-577
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• 2009

Dynamic behavior simulation of supersonic engine was performed and PI control algorithm was studied for the buzz control in the inlet and the thrust control. Firstly, required thrust was tracked according to the fuel flow control and then inlet pressure was regulated through the nozzle throat area control so that the buzz margin has the positive all the time. The control was performed according to the change of flight Mach number, altitude and angle of attack. The proportional gain and the integral gain for regulating the buzz margin was induced and simulated. In the results, it was confirmed and satisfied that control target in the operating area was changed the angle of attack from $0^{\circ}$ to $10^{\circ}$ at the flight Mach number of 2.1~3.0.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.4
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• pp.33-38
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• 2003

The Physics of the Plume-induced shock and separation Particularly at a high Plume to exit pressure ratio and supersonic speeds up to Mach 3.0 with and without a passive control method, porous extension, were studied using computational techniques. Mass-averaged Navier-Stokes equations with the RNG $\kappa$-$\varepsilon$ turbulence model were solved using a fully implicit finite volume scheme and a 4-stage Runge-Kutta method. The control methodology for plume-afterbody interactions is to use a perforated wall attached at either the nozzle exit or the edge of the missile base. The Effect of porous wall length on plume interference is also investigated The computational results show the main effect of the porous extension on plume-afterbody interactions is to restrain the plume from strongly underexpanding during a change in flight conditions. With control, a change in porous extension length has no significant effect rut plume interference.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.97-102
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• 2013

A prestudy on expendable turbine engine for high-speed vehicle was conducted. After two possible mission profiles were established to decide the engine requirements, design point analysis was performed with the values of design parameter which were obtained from similar class engines, references, etc. The results showed that specific net thrust and specific fuel consumption with turbine inlet temperature of 3,600 R are 2,599.4 ft/s and 1.483 lb/(lb*h) respectively at the flight condition of sea level, Mach 1.2. It was also found that major design parameters for determining maximum net thrust were turbine inlet temperature for low supersonic and transonic flight speed and compressor exit temperature for high supersonic flight speed from the results of performance analysis on the two possible mission profiles. In addition, simple turbojet engine with an axial compressor, a straight annular combustor, an one stage axial turbine and a fixed throat area converge-diverge exhaust nozzle was proposed as the configuration of simple low cost lightweight turbine engine.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.1
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• pp.12-20
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• 2002

An integrated analysis of kerosine/LOX based KSR-III rocket body/plume flowfield has been performed. The analysis has been executed employing three kind of gas thermo-chemical models including calorically perfect gas, multiple species chemically reacting gas, and chemically frozen gas models and their effect on rocket flowfield has been accessed to provide the most appropriate gas thermo-chemical model which meets a specific purpose of performing rocket body and plume analysis. The finite-rate chemically reacting flow solution exhibited higher temperature throughout the flowfield than other gas models due to the increased combustion gas temperature caused by the chemical reactions within the nozzle. All the reactions were dominated only in the shear layer and behind the barrel shock reflection region where the gas temperature is high and the effect of finite-rate chemical reactions on the flowfield was found to be minor. However, the present plume computation including finite-rate chemical reactions revealed major reactions occurring in the plume and their reaction mechanisms and as well.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.604-605
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• 2017

The present work aims at visualization of the supersonic air jet flows discharged from C-D nozzles. In the present experiments, Prticle Image Velocimetry (PIV) was employed to specify the jet flow field quantitatively, and a color Schlieren optical method was applied to observe the same jets qualitatively. The $0.5{\mu}s$ duration of spark light source was used for Schlieren and it can be controled as $0.5{\mu}s$, $1{\mu}s$, $2{\mu}s$ and focusing mode. The convergent-divergent nozzles were used to generate the jet flow with the design Mach number of 2.0, 2.2. Nozzle pressure ratios (NPRs) were varied from 5 to 8. A good comparison of the jet size and shock location from the Schlieren images with the PIV quantitative values is obtained. The obtained images clearly showed the major features of the under-expanded jet, over-expanded jet, sound wave, turbulent eddies and so on.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.41 no.9
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• pp.563-570
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• 2017

Experimental and numerical studies were performed to investigate the performance and internal flow characteristics of a supersonic second throat exhaust diffuser (STED) with back pressure ($P_a$). An ejector system was used to vary the back pressure ($P_a$) conditions. The operating gas for the STED and the ejector was high pressure nitrogen at room temperature. When the back pressure ($P_a$) at a constant nozzle inlet pressure $P_0$) decreases, the pressure recovery location moves downstream. If the pressure ratio $P_0/P_a$) is the same, even if the nozzle inlet pressures $P_0$) are different, the diffuser's internal flow pattern and starting pressure ratio ($(P_0/P_a)_{st}$) are almost the same.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.680-685
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• 2008

The ejector is a simple device which can transport a low-pressure secondary flow by using a high-pressure primary flow. In general, it consists of a primary driving nozzle, a mixing section, and a diffuser. The ejector system entrains the secondary flow through a shear action generated by the primary jet. Until now, a large number of researches have been made to design and evaluate the ejector systems, where it is assumed that the ejector system has an infinite secondary chamber which can supply mass infinitely. However, in almost all of the practical applications, the ejector system has a finite secondary chamber implying steady flow can be possible only after the flow inside ejector has reached an equilibrium state after the starting process. To the authors' best knowledge, there are no reports on the starting characteristics of the ejector systems and none of the works to date discloses the detailed flow process until the secondary chamber flow reaches an equilibrium state. The objective of the present study is to investigate the starting process of an ejector-diffuser system. The present study is also planned to identify the operating range of ejector-diffuser systems where the steady flow assumption can be applied without uncertainty. The results obtained show that the one and only condition in which an infinite mass entrainment is possible is the generation of a recirculation zone near the primary nozzle exit. The flow in the secondary chamber attains a state of dynamic equilibrium at this point.

• Journal of Mechanical Science and Technology
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• v.16 no.11
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• pp.1448-1456
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• 2002

The present study addresses the flow characteristics involved in the self-induced oscillations of the underexpanded jet impinging upon a cylindrical body. Both experiment and computational analysis are carried out to elucidate the shock motions of the self-induced oscillations and to find the associated major flow factors. The underexpanded sonic jet is made from a nozzle and a cylindrical body is placed downstream to simulate the impinging jet upon an obstacle. The computational analysis using TVD scheme is applied to solve the axisymmetric, unsteady, inviscid governing equations. A Schlieren system is employed to visualize the self-induced oscillations generated in flow field. The data of the shock motions are obtained from a high-speed video system. The detailed characteristics of the Mach disk oscillations and the resulting pressure variations are expatiated using the time dependent data of the Mach disk positions. The mechanisms of the self-induced oscillations are discussed in details based upon the experimental and computational results.