• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.31-36
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• 2015

As one of the most promising propulsive systems in the future, the dual-mode scramjet engine has drawn the attention of many researches. Detailed flow features concerned with the isolator play an important role in the dual-mode scramjet system. The 2D numerical method has been used for the dual-mode scramjet with wind tunnel. To validate the ability of the numerical model, numerical results have been compared with the experimental results. Overall pressure distributions show quite good match with the experimental results. Back pressure has been studied for maximum pressure rising. According to the results, pressure distribution of supersonic inlet section is not influenced by back pressure. The shock train is pushed towards upstream as the back pressure increases. The maximum value of back pressure without inlet unstart goes up rapidly and then keeps constant when the isolator length increases. The optimal length of isolator section ($L/H_{th}$) is 8.7 in this model.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.2
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• pp.10-18
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• 2010

Scramjet engine is one of the core parts of hypersonic vehicle of next generation and being investigated by many countries. Korea Aerospace Research Institute performed a ground test of the model scramjet engine S1 in 2007. And, S2 model which is improved from S1 model in engine startability and thrust was tested with HIEST (High Enthalpy Shock Tunnel) at Kakuda Space Center of JAXA. Design condition of S2 model was Mach 6.7, however, it was tested at Mach 7.7 as an off-design condition test. As a test result, flow separation was found at the inside of the intake, but the engine showed stable combustion pressure distribution. Furthermore, compared to other test models, S2 model showed a good performance value in thrust and specific impulse.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.11
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• pp.1067-1074
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• 2010

In this study, RBDO(Reliability Based Design Optimization) was performed for a supersonic double-wedge inlet. By considering uncertainty of design with given design space, the pressure recovery was transformed into the probabilistic constraint while the inlet drag was considered as a deterministic objective function. To save computational analysis cost and to search good design space, Latin-Hypercube design of experiment and the Kriging model were incorporated and then RBDO was performed. Monte-Carlo simulation was performed to verify the accuracy of AFORM(Advanced First Order Reliability Method). It was found that AFORM result agreed very well with the Monte-Carlo simulation result. The system reliability was guaranteed by considering uncertainty of the design variables. In case of considering diverse uncertainty of system design, RBDO was found to be useful.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.7
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• pp.529-537
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• 2021

Wind tunnel tests were carried out with a scramjet high speed vehicle. Since the scramjet engine does not have a compressor, it has a simple structure, but it is important to design the intake for the supersonic combustion in the combustion chamber. In this study, internal flow characteristics and the starting condition were analyzed by measuring the pressure at the isolator exit just before the combustion chamber, and the intake performance parameters were calculated and compared the result on every Mach number. The aerodynamic characteristics of the flow-through high speed vehicle were analyzed and internal drag correction is required to precisely analyze the aerodynamic characteristics. In this paper, an experimental technique using three probes for internal drag correction was proposed. By applying internal drag correction, it was able to figure out the effect of the internal flow on the aerodynamic force of the vehicle.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.5
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• pp.60-65
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• 2011

The development of an air data acquisition algorithm has been described in the supersonic flow at the preliminary design stage with pressure data acquisition device composed of major three total pressure sensors and two static pressure sensors which are installed on the surface of a cone type supersonic inlet. Through this algorithm, Mach number, angle of attack and sideslip angle can be very easily derived with simple interpolation algorithm and predefined data tables. The available range of Mach number is 1.6 to 4.0, angle of attack, $-12^{\circ}$ to $12^{\circ}$ and sideslip angle, $-12^{\circ}$ to $12^{\circ}$. In preliminary design stage, the data tables applied to the developed algorithm are constructed with data driven by Taylor Maccoll equation. The present algorithm would be useful to get supersonic flow air data for the various aerial vehicles and their flight tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.14-14
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• 1998

램제트 엔진은 초음속에서 공기가 충격파를 통해 아음속으로 속도가 낮아지고 압력이 증가하는 램 압축 현상을 이용하되 압축기를 사용하지 않고 아음속 상태에서 연소하는 구조로 되어 있다. 따라서 각 부품의 성능은 독립적이지 않으며 전체적인 성능을 규명하기 위해서는 공기 흡입구와 연소실, 연료 분사체계 등의 상호작용을 고려하여 유동의 특성과 그에 따른 연소현상의 특성을 함께 고려해야만 한다. 본 연구에서는 속도 범위 Mach 5이내, 고도 30km이내의 운항조건을 갖는 유도 무기에 장착될 램제트 엔진 개발을 위한 연구의 첫 번째 단계로써, 램제트 연소실내의 유동장해석을 실험과 수치해석 두 분야로 나누어 수행하였다.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.14 no.2
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• pp.27-32
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• 2006

The performance-design algorithm of a liquid ramjet engine was studied, which showed the comparable results with CFD calculation except the shock-boundary layer interaction region. In addition to the description of the design algorithm, several important design parameters, such as equivalence ratio, combustion efficiency, air capturing factor, and flight Mach number, are evaluated as predicting one of performance factors, Isp, of a typical ramjet engine, so that the flight envelope might not be determined with narrow perspective performance-operation-area in off-design regime.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.5
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• pp.65-76
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• 2020

On the initial design process of a scramjet vehicle such as the trajectory prediction, it is inevitable to estimate the aerodynamic performance of a three-dimensional effect. Despite the necessity of intensive computing for the three-dimensional model, it is inefficient in predicting a wide range of aerodynamic performance. In this study, an engineering model for aerodynamic performance was developed based on two-dimensional computational fluid analysis and linearized supersonic inviscid flow theory. Correspondingly, the three-dimension aerodynamic performance relations are presented based on the two-dimensional results. And the additional three-dimensional computation was performed to evaluate the adequacy for the extended relations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.10-10
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• 2000

램제트 엔진은 비추력이 높고 추력 레벨은 낮으므로, 2단 추진기관에 적합한 추진 시스템이다. 1단-추진기관의 작동이 끝나고, 2단 램제트 엔진이 점화 후 안정된 연소에 도달되기까지 비행체의 속도는 항력에 의하여, 초당 약 마하수 0.1 정도씩 감소된다. 1단 연소 후 2단 램제트로 전환되는 지연시간이 길수록 1단에서 요구되는 종말 가속도는 증가되므로, 1단이 차지하게되는 부피는 증가되고 비행체의 크기 또한 늘어나게 된다. 따라서 1단에서 2단 램제트로 천이되는데 소요되는 시간을 가능한 짧게 하는 것이 효과적이다. 그러나 램제트 엔진의 특성상 선결되어야할 다음과 같은 여러 문제들이 있다. 첫째, 1단 작동 시 공기 흡입구와 연소실은 차단벽으로 분리되어 있다가, 1단 연소후 차단막이 제거되어 외부공기가 램제트 연소실로 흡입된다. 흡입되는 공기는 흡입구의 형상에 의하여 램 압축되지만 초음속으로 연소실을 통과하게된다. 연료 주입 구에서 공급되는 연료는 연소실에서 유동의 흐름방향(streamline)에 따라서 연소실로 확산되는데, 연소되기 전에는 유속이 빠르게 노즐로 빠져 나가므로 램제트 연료가 재순환 구역(recirculation zone)으로 침투하는데 쉽지가 않다. 둘째, 연소실 입구에서 발생되는 와류 (ring vortex)는 1단 연료의 고온 연소 가스를 연소실로 확산시키는데, 비 균일한 온도 분포를 유발하여 램제트 연료의 점화에너지가 공급되는 시간이 적당하지 않을 경우 균일한 화염 전파에 악영향을 준다. 셋째, 연소실에서의 빠른 유동 조건은 연료가 연소실에 머무를 수 있는 시간을 감소시키며, 연소실 입구에서 강한 전단 응력이 발생되어 화염이 안정화되는데 악 영향을 미치게된다. 본 논문은 공기 흡입구, 연소실 및 노즐을 통합하여 수치해석을 하였으며 열유동/점화/연소등의 미케니즘을 이해하고, 주요 인자들 중 와류의 영향에 초점을 맞추었다.다고 판단되며 배기 가스 자체에 대기 공기중에 함유되어 있던 습기가 얼어붙는(Icing화) 문제가 발생하기 때문에 배기가스의 Icing을 방지하기 위하여 압축기 끝단에서 공기를 추출하여 배기부분에 송출할 필요성이 있는 것으로 판단되었다. 출구가스의 기체 유동속도가 매우 빠르므로 (100-l10m.sec) 이를 완화하기 위한 디퓨저의 설계가 요구된다고 판단된다. 또 연소기 후방에 물을 주입하는 경우 열교환기 및 기타 부분품에 발생할 수 있는 부식 및 열교환 효율 저하도 간과할 수 없는 문제로 파악되었다. 이러한 기술적 문제가 적절히 해결되는 경우 비활성 가스 제너레이터는 민수용으로는 대형 빌딩, 산림, 유조선 등의 화재에 매우 적절히 사용되어 질 수 있을 뿐 아니라 군사적으로도 군사작전 중 및 공군 기지의 화재 그리고 지하벙커에 설치되어 있는 고급 첨단 군사 장비 등의 화재 뿐 아니라 대간첩작전 등에 효과적으로 활용될 수 있을 것으로 판단된다.가 작으며, 본 연소관에 충전된 RDX/AP계 추진제의 경우 추진제의 습기투과에 의한 추진제 물성 변화는 미미한 것으로 나타났다.의 향상으로, 음성개선에 효과적이라고 사료되었으며, 이 방법이 편측 성대마비 환자의 효과적인 음성개선의 치료방법의 하나로 응용될 수 있으리라 생각된다..7%), 혈액투석, 식도부분절제술 및 위루술·위회장문합술을 시행한 경우가 각 1례(2.9%)씩이었다. 13) 심각한 합병증은 9례(26.5%)에서 보였는데 그중 식도협착증이 6례(17.6%), 급성신부전증 1례(2.9%), 종격동기흉과 폐염이 병발한 경우와 폐염이 각 1례(2.9%)였다. 14) 식도경 시행회수는 1회가 17례(54.8%), 2회가 9례(29.0%), 3회 이상이 5례(16.1%)였다.EX>$IC_{50}$/ 값이 210 $\mu\textrm{g}$/$m\ell$로서 효과적

• Journal of the Korea Academia-Industrial cooperation Society
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• v.21 no.8
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• pp.215-221
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• 2020

Navigation lights used in supersonic aircraft are used to identify the direction and location of the aircraft. The color of the navigation lights and location of installation are defined by aviation law as red for the left wing, green for the right wing, and white for the tail. Navigation lights operate in BRT and DIM modes. BRT is the brightest mode, and DIM is an output with dimmed brightness. Navigation lights serve to prevent aircraft collisions and are very important for stability and location identification. One phenomenon is that the inlet and tail navigation lights flicker abnormally. In this study, fault tree analysis was performed in two stages. The first step was derived from three causal factors, the second step developed five improvements, and the optimal improvement plan was drawn. The navigation lights confirmed that the initial input power was unstable as the main cause of abnormal flickering. As an improved method, the circuit was adjusted to stabilize the initial power, and it was confirmed that flickering did not occur as a result of the tests under the same conditions.