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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.
한국항공우주연구원 가스터빈엔진 고고도 성능시험설비 소개
이경재(Kyungjae Lee),양인영(Inyoung Yang),김춘택(Chuntaek Kim),양수석(Sooseok Yang) 한국추진공학회 2014 한국추진공학회 학술대회논문집 Vol.2014 No.12
한국항공우주연구원은 소형 터보제트/팬 엔진의 고고도 성능시험이 가능한 설비를 1999년 국내 최초로 설치 완료하여, 현재까지 운용 중에 있으며, 2003년에는 KOLAS(공인시험기관인정기구)로부터 국제공인시험기관 인정을 획득하여 설비의 국제적인 신뢰도를 확보하였다. 2008년에는 터보샤프트엔진의 시험이 가능하도록 설비를 업그레이드 하여 수리온 엔진의 고고도 성능시험을 성공적으로 완료하였다. 이 외에도 설비의 저온모사 설비를 활용하여 국내개발 전압관의 방빙성능 테스트를 수행완료 하였으며, 고압공기를 활용한 RBCC(Rocket Based Combined Cycle) 엔진의 이젝터 성능시험도 수행하는 등 다양한 영역에서 설비가 활용되고 있다. 본 설비에 장착되어 있는 총 3대의 대유량 압축기의 구성을 변경하여 최고 40,000 ft의 고도와 마하수 1까지 모사가 가능하며, 최대 입구공기 유량과 압력은 각각 40㎏/s@2bar와 4bar 이다. An altitude test facility(AETF) for gas turbine engines was built at the Korea Aerospace Research Institute(KARI) in 1999 and has been being operated well until now. AETF acquire international certification from KOLAS(Korea Laboratory Accreditation Scheme) about thrust and SFC of turbo-jet/fan engine for international reliability. At 2008, an AETF of KARI has capability of turbo-shaft engine altitude test with upgrade of facility and altitude engine test of SURION was performed successfully. Above this, AETF of KARI has been used in the various area of aerospace. Anti-icing test of total pressure probe and ejector mode test of RBCC(Rocket Based Combined Cycle) engine also performed with low-temp and high pressure air source capacity, respectively. AETF can simulate up to 40,000 ft and Mach 1 with three large scale compressor. Maximum inlet air flow and pressure is 40㎏/s@2bar and 4bar@15㎏/s.
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.