Ki-Young Hwang

Effects of density change and natural convection on the solidification process of a pure metal

Abstract The solidification of an initially superheated metal within a rectangular enclosure is studied numerically through the front-fixing method accounting for the solid-liquid density change and natural convection in the melt. The investigation focuses on the influence of both the volume contraction caused by solid-liquid density change and the natural convection effect in the melt. Difficulties associated with the complex time-dependent solid and liquid domains, the shapes of which are also a part of the solutions, are overcome by employing the boundary-fitted coordinate system. The computed results are validated and are shown in the forms of the transient position of the interface, the temperature distribution, the flow pattern, the Nusselt number and the solidification fraction, with respect to time.

Porous silica aerogel/honeycomb ceramic composites fabricated by an ultrasound stimulation process

The synthesis behavior of nanoporous hydrophobic silica aerogel in honeycomb-type ceramics was observed using TEOS and MTES. Silica aerogel in the honeycomb ceramic structure was synthesized under ultrasound stimulation. The synthesized aerogel/honeycomb ceramic composites were dried under supercritical CO2 drying conditions. The values for the line shrinkage of the wet gels during supercritical CO2 drying declined from 19% to 4% with an increase in the H2O/TEOS molar ratio from 8 to 24. Low shrinkage was a key factor in increasing the interface compatibility with the aerogel/honeycomb ceramic composites. The optimum condition of silica aerogel in the honeycomb-type ceramic structure had a TEOS:MTES: H2O:glycerol ratio equal to 1:1.2:24:0.05 (mol%).

A Study on a Modeling and Simulation Program of an Environmental Control System with a Phase Change Heat Exchanger

A study on the modeling and simulation of an environmental control system (ECS) is described in this paper. While a conventional ECS mainly consisted of an air cycle machine and heat exchangers, a new concept of a phase change heat exchanger was added to improve the transient performance of the ECS. As a result, an ECS modeling program including the phase change heat exchanger is newly developed to estimate its effect in various flight conditions such as take-off, maneuver, cruise, and landing. The simulation result regarding a virtual flight profile has confirmed the new ECS fulfilled the requirement by showing the temperature of the cooling air returned from the bay was always kept below 80℃. Through this study, the new ECS concept with PCHE was verified successfully.

A Study on Injection Characteristics of High Temperature Fuel through Orifice Injectors

An experimental study was conducted to study fuel injection characteristics through plain orifice injectors when the fuel was heated to the temperature higher than its boiling point. Three injectors with different orifice diameters were used to measure the flow coefficient (α) for the injection pressure ranges of 3, 5, and 10 bar and the fuel temperature ranges between 50 and 270°C. The study showed that α decreases gradually with the fuel temperature below 180°C while it drops abruptly when the temperature goes beyond 187°C, the boiling temperature of the fuel. The slope of α bifurcated at the boiling temperature for different injection pressures, and α decreased faster for the lower injection pressure due to the more active boiling in the injector. In addition, the larger orifice diameter had the higher α value, and α jumped at moderate temperature ranges when the injection pressure was low, implying the turbulent-laminar transition phenomena. The measured α was plotted against the cavitation number(Kc), and the characteristics were independent of the applied pressure for small injectors when the fuel was evaporated before it was injected.

A Study on High-Temperature Fuel Injection Characteristics through Swirl Injectors

An experimental study was conducted to investigate fuel injection characteristics through swirl injectors when the fuel was heated to very high temperature conditions. Three swirl injectors with different orifice diameters and swirler geometries were used to measure the flow coefficient () for the injection pressure ranges between 3 and 10 bar and the fuel temperature from 50 to . The results showed that the variation characteristics of with respect to cavitation number () were highly dependent on both the orifice diameter and the swirler geometry. In addition, the characteristics of variation with respect to AR, the area ratio of the flow through the swirler and the orifice, has revealed that the effect of boiling is retarded but the slope of decreasing after the boiling effect is present tumbles as AR increases.

Research Activities of Transpiration Cooling for High-Performance Flight Engines

Transpiration cooling is the most effective cooling technique for the high-performance liquid rockets and air-breathing engines operating in aggressive environments with higher pressures and temperatures. When applying transpiration cooling, combustor liners and turbine blades/vanes are cooled by the coolant(air or fuel) passing through their porous walls and also the exit coolant acting as an insulating film. Practical implementation of the cooling technique has been hampered by the limitations of available porous materials. But advances in metal-joining techniques have led to the development of multi-laminate porous structures such as Lamilloy fabricated from several diffusion-bonded, etched metal thin sheets. And also with the availability of lightweight, ceramic matrix composites(CMC), transpiration cooling now seems to be a promising technique for high-performance engine cooling. This paper reviews recent research activities of transpiration cooling and its applications to gas turbines, liquid rockets, and the engines for hypersonic vehicles.

Local Heat and Mass Transfer Characteristics for Multi-Layered Impingement/Effusion Cooling

Multi-layered impingement/effusion cooling is an advanced cooling configuration that combines impingement jet cooling, pin cooling, and effusion cooling. The arrangement of the pins is a critical design factor because of the complex heat transfer in the internal structure. Therefore, it is important to measure the local heat transfer at all internal surfaces as a function of the pin spacing. In this study, a naphthalene sublimation method was employed to measure the details of the heat/mass transfer at the internal surfaces, including the injection plate, effusion plates, and the pins. An staggered array of holes was formed at the injection plate and effusion plates where the ratio of the height to the diameter of the pins, h/d, was fixed at 0.25. The ratio of the pin spacing to the diameter, sp/d, was varied in the range 1.5≤sp/d≤6, and the Reynolds number based on the hole diameter was 3000. As a result, a vortex ring formed near the pin, leading to re-impingement flows in the narrow channel. The jet flow impinged strongly on the pin, resulting in a large heat transfer region at each surface. The total average Sherwood number with sp/d=1.5 was larger than that with sp/d=6 by a factor of 1.5.Copyright

Prestudy on Expendable Turbine Engine for High-Speed Vehicle

․ KiYoung Hwang*ABSTRACT A prestudy on expendable turbine engine for high-speed vehicle was conducted. After two possible mission profiles were established to decide the engine requirem ents, 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 an d 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 parame ters 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 throa t area converge-diverge exhaust nozzle was proposed as the configuration of simple low cost lightweight turbine engine.초 록 초고속 비행체에 적용 가능한 소모성 터빈엔진 개발을 위한 사전연구를 수행하였다. 엔진 요구도 결정을 위한 가상 운용임무형상을 선정한 후, 유사급 엔진과 참고문헌 등을 통해 확보된 설계변수 값을 활용하여 설계점 해석을 수행하였는데, 해면고도, 마하수 1.2 조건에서 터빈입구온도 3,600 R에 대한 설계점 계산결과, 비추력 2,599.4 ft/s, 비연료소모율 1.483 lb/(lb*h)이 예측되었다. 두 가지 임무형상에 대한 엔진 성능해석결과로부터 엔진 최대 순추력을 결정하는 설계변수는 천음속 및 낮은 초음속영역에서는 터빈입구온도, 높은 초음속 영역에서는 압축기 출구온도임을 확인하였다. 이밖에도 단순, 저가, 경량의 터빈엔진형상으로 축류형 다단압축기와 직류형 연소기, 1단 축류터빈, 고정 수축팽창 노즐이 적용된 단순터보제트엔진을 제시하였다. Key Words: Expendable Turbine Engine(소모성 터빈엔진), High-speed Vehicle(초고속 비행체), Mission Profile(임무형상), Design Point(설계점), Turbine Inlet Temperature(터빈입구온도)

Optimization of Turbofan Engine Design Point by using Seven Level Orthogonal Array

aerokim@dreamwiz.comABSTRACT For design optimization, engineers should require the accurat e information of design space and then explore the design space and carry out optimization. Recently, the total design framework, based on design of experiments and optimization, is widely used in indus try areas to explore the design space above all. For optimizing turbofan engine design point, the res ponse surface model is constructed by using the 7 level orthogonal array which satisfies the statisti cal uniformity and orthogonality and gets the dense design space information. The multi-objective genetic algorithm is used to find the optimal solution within the given constraints for finding global optima l one in response surface model. The optimal solution from response surface model is verified with G asTurb simulation result.초 록 설계 최적화를 위해서 설계자는 우선적으로 설계영역 전반에 걸쳐 정확한 정보를 획득하고, 설계영역 탐색을 실시한 후에 최적화를 실시해야 한다. 최근에 설계영역 탐색을 우선적으로 실행하기 위하여 실험계획법과 반응표면모델에 최적화를 적용하는 통합설계 프레임워크의 적용이 산업체 전반으로 일반화되고 있다. 본 연구에서는 터보팬 엔진 설계점 최적화를 위하여, 설계영역의 정보를 조밀하게 얻을 수 있으며 통계학적인 직교성과 균형성을 모두 만족하는 7 수준 직교배열을 생성한 후에 반응표면모델을 구축한다. 구축된 반응표면모델에 전역 최적값을 찾기 위하여 다목적 유전자알고리즘을 적용하여 주어진 제한조건을 만족하는 최적값을 찾아 GasTurb 결과와 검증을 수행한다.Key Words: Turbofan(터보팬), Orthogonal Array( 직교배열), Design of Experiments(실험계획), Response Surface Model(반응표면모델), Optimization(최적화)Received 27 November 2012 / Revised 18 March 2013 / Accepted 25 March 2013

Research Activities about Characteristics of Fuel Injection and Combustion Using Endothermic Fuel

5genie@hanmail.netABSTRACT Endothermic fuel utilizing technology is considered as a uniq ue practical method of hypersonic vehicle for long distance flight. Research activities about cha racteristics of fuel injection and combustion using cracked by endothermic reaction are reviewed. Studies on characterization of supercritical fuel injection and mixing within supersonic flow field are surveyed. Researches on combustion characteristics such as ignition delay time, laminar burning ve locity and combustion efficiency at supersonic model combustor are reviewed. In addition, domestic research activities on endothermic fuel are surveyed.초 록 장거리 극초음속 비행체에 적용 가능한 유일한 냉각방안으로 알려져 있는 흡열연료 적용기술을 개발하기 위하여 흡열반응에 의해 분해된 연료의 분사 및 연소특성에 대한 연구사례를 살펴보았다. 흡열반응을 거친 연료가 연소실에 분사될 때 처해지는 초임계 상태의 분사 특성, 초임계 연료가 초음속 유동장에 분사될 때의 공기혼합 특성 등에 관한 연구사례를 살펴보았고, 연소특성으로서 점화지연시간 및 화염전파 속도에 미치는 영향, 초음속 연소실에서 연소될 때의 연소효율 상승 연구사례 등을 살펴보았다. 국내에서 수행된 흡열연료 관련 연구동향을 살펴보았다.Key Words: Endothermic Fuel(흡열연료), Hypersonic Propulsion( 극초음속 추진), Supercritical Injection(초임계 연료분사), Combustion Characteristics(연소 특성)Received 30 November 2012 / Revised 1 April 2013 / Accepted 11 April 2013