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차봉준,임병준,양수석,이대성,Cha, Bongjun,Lim, Byungjun,Yang, Sooseok,Lee, Daesung 한국유체기계학회 2001 한국유체기계학회 논문집 Vol.4 No.1
The experimental study on the effect of axial clearance between the tip of impeller blades and stationary shroud has been performed. The investigated compressor, which is a part of a small auxiliary power unit engine, consists of a curved inlet, a centrifugal impeller, a channel diffuser and a plenum chamber. It was designed for a total pressure ratio of 4.3 and an efficiency of $77\%$ at design speed of 60,000 rpm. The experiments are carried out in an open-loop centrifugal compressor test rig driven by a turbine. For the four different clearance ratios Cr(clearance/impeller tip width) of 6.25, 10.93, 15.60 and 20.30 percent, the overall performance data are obtained at $97\%,\;90\%$ and $80\%$ of the design speed. The results show the overall pressure ratio decrease of $7.7\%$ and the efficiency loss of $8.7\%$ across the variation of clearance ratio near the design speed. It also indicates that the influence of tip clearance became weaker as the flow rate is reduced and the stable operating range is not significantly influenced by the change of clearance ratio.
김진한,최창호,김춘택,양수석,이대성,Kim, Jin-Han,Choi, Chang-Ho,Kim, Chul-Taek,Yang, Sooseok,Lee, Daesung 한국유체기계학회 2001 한국유체기계학회 논문집 Vol.4 No.1
In developing a multistage compressor, the stage matching is one of the critical design issues. The mismatching can be often observed even if each stage has been proven good and then used as part of a compression system. A good matching among the stages can be achieved by changing various design parameters (i.e., passage cross sectional areas, blades angles, stagger angles, curvature, solidity, etc.). Therefore, designers need to find out what parameters must be changed and how much. In this study, a method to search the design parameters for optimum stage matching has been used based on an 1-D mathematical model of a compressor, which uses the data obtained from the preliminary test to identify the design parameters. This methodology is applied with a two-stage axial compressor, which was originally designed for a helicopter gas turbine engine. After identifying design parameters using preliminary test data, an optimization process has been employed to achieve the best matching between the stages (i.e., maximum efficiency of the compressor at its operation modes within a given range of the rotor speed under given restrictions for required stall margins and mass flow). 3-D flow calculations have been performed to confirm the usefulness of the corrections based on the 1-D mathematical model. Calculational results agree well with the experimental data in view of the performance characteristics. Some promising results were produced through the methodology proposed in this paper in conjunction with flow calculations.
조남경(Namkyung Cho),양수석(Sooseok Yang) 한국추진공학회 2023 한국추진공학회지 Vol.27 No.5
Nuclear space propulsion has the advantage of superior specific impulse compared to the currently widely used chemical propulsion, the system can be simplified by omitting the oxidizer system, and has the ability to improve performance in the future. The high specific impulse makes missions that are only possible with large chemical propulsion launch vehicles possible with medium-sized launch vehicles. This paper describes the characteristics of nuclear propulsion that are different from chemical propulsion and introduces the nuclear propulsion system and key components. Nuclear thermal propulsion and electric propulsion, and a bimodal system and a hybrid system that combines chemical propulsion and nuclear propulsion are introduced. Also, the development aspects of nuclear propulsion as well as test facility and launch safety of the nuclear power propulsion system are discussed.
이경재(Kyungjae Lee),강상훈(Sanghun Kang),양수석(Sooseok Yang),박철(Chul Park) 한국추진공학회 2011 한국추진공학회 학술대회논문집 Vol.2011 No.5
비행 마하수 6으로 운용되며, 지상 정지 추력으로부터 사용이 가능하도록 2단 추진체 개념이 적용된 스크램제트 엔진 비행체에 대한 개념설계를 수행하였다. 1단은 로켓을 적용되었으며, 2단은 탄화수소 계열의 연료를 사용하는 스크램제트 엔진이 적용되었다. 개념설계를 위하여 반경 2,000km의 운용거리와 0.2톤의 탑재체 무게를 가정하였다. 개념설계의 첫 번째 단계로 3-DOF 코드를 이용하여 비행궤도를 계산하였으며, 계산된 비행계도를 바탕으로 일차원-비평형 유동 코드와 NASA의 HASA 데이터베이스를 이용하여 스크램제트 엔진에 대한 개념설계를 수행하였다. In this study, two-stage hypersonic scramjet vehicle was designed for the flight condition of Mach number 6. In order to launch at sea level and Mach number 0, two stage concept was applied. The first stage of the vehicle is rocket-powered and is mounted under the second stage. The second stage is scramjet-powered propulsion system and has wing. The suggested mission scenario is to deliver 0.2 ton payload to the range less of 2000km. For the first step of conceptual design, trajectory of air vehicle was calculated by 3-DOF trajectory code. Based on the result of trajectory code, scramjet engine design and mass estimation were performed by non-equilibrium nozzle flow code and NASA"s HASA model, respectively. In order to find best solution, all step of designing process was iterated until they were converged.