• 한국항공우주학회지
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• 제50권1호
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• pp.21-30
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• 2022

스크램제트 엔진은 운용 중에 비시동이 발생하면 재시동이 거의 불가능하다. 그러므로 이에 대한 예측이 매우 중요하다. 본 연구에서는 격리부 출구에서의 배압을 조절함에 따라 나타나는 비시동 과정을 수치적으로 모사하였다. 비시동 데이터 검출은 벽면에서의 압력 데이터에 서포트 벡터 머신(SVM) 기법을 적용하여 흡입구의 시동과 비시동 데이터로 분류하였고, 시동과 비시동의 분류에 가장 적합한 압력 센서의 위치를 도출하였다. 또한 엔진의 시동과 비시동 경계를 분석하여 엔진이 비시동 되기까지의 여유(마진)을 정량화하였다.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2012년도 제38회 춘계학술대회논문집
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• pp.388-391
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• 2012

마하 5 스크램젯 엔진에 대하여 흡입구 시동 특성 시험을 수행하였다. 시험 모델과 시험 설비의 상호 작용이나 과다한 내부 압축비에 의한 흡입구 불시동 현상을 관찰하였다. 모델의 위치를 조정함으로써 모델-설비의 상호 작용을 없애고 카울의 형상을 변경하여 내부 압축비를 조정함으로써 흡입구 시동을 달성하였다.

• Advances in aircraft and spacecraft science
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• 제5권4호
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• pp.431-444
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• 2018

Numerical simulations have been conducted to study the unstart/restart characteristics of an over-under turbine-based combined-cycle propulsion system (TBCC) inlet during the inlet transition phase. A dual-solution area exists according to the Kantrowitz theory, in which the inlet states may be different even with the same input parameters. The entire transition process was divided into five stages and the unstart/restart hysteresis loop for each stage was also obtained. These loops construct a hysteresis surface which separates the operating space of the engine into three parts: in which a) inlet can maintain a started state; b) inlet keeps an unstarted state; c) inlet state depends on its initial state. During the transition, the operation of the engine follows a certain order with different backpressures and splitter angles, namely control route, which may result in disparate inlet states. Nine control routes with different backpressures and transition stages were designed to illuminate the route-dependent behavior of the inlet. The control routes operating towards the unstart boundary can make the inlet transit from a started state into an unstarted one. But operating backward the same route cannot make the inlet restart, additional effort should be made.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2004년도 제22회 춘계학술대회논문집
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• pp.26-33
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• 2004

In this study, a boundary-layer bleeding and a two-staged fuel injection were applied to a scramjet engine for suppressing unstart transition and improving the thrust performance under Mach 6 flight conditions. With the boundary-layer bleeding, the engine could operate without unstart transition around at the fuel equivalence ratio of unity ($\Phi$ = 1). The thrust increment from the no fuel condition (dF) increased to 2460 N, which was about 1.4 times as large as that of the case without the bleeding and maximum in our Mach 6 tests. It was confirmed that the boundary-layer bleeding suppressed the separation during the engine operation. The two-staged fuel injection was less effective for improving the thrust performance com-pared with the single-staged one with the bleeding at Mach 6.

• 대한기계학회논문집B
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• 제26권6호
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• pp.765-771
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• 2002

This Paper presents numerical solutions of two-dimensional Euler equations for supersonic steady and unsteady flows with heat addition in a convergent-divergent duct, The Van Leer FVS (flux vector splitting) method in generalized coordinates is employed in order to calculate the inviscid strong shock waves caused by thermal choking. We discuss on transient characteristics, start and unstart phenomena caused by thermal choking, limit of equivalence ratio to avoid thermal choking and fluctuation of specific thrust caused by thermal choking. We prove that thermal choking is a serious problem in view of engine performance.

• 대한기계학회논문집B
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• 제29권1호
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• pp.80-86
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• 2005

This study presents numerical solutions of the two-dimensional Navier-Stokes equations for supersonic unsteady flow in a convergent-divergent nozzle with heat addition. The TVD scheme in generalized coordinates is employed in order to calculate the moving shock waves caused by thermal choking. We discuss on transient characteristics, start and unstart phenomena, fluctuations of specific thrust caused by thermal choking and viscous effects. We prove that the control of separation of boundary layer is the most important key problem to prevent the thermal choking.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2004년도 제22회 춘계학술대회논문집
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• pp.649-656
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• 2004

Hypersonic wind tunnel test of the rectangular variable geometry intake is performed. For realization of a Precooled turbojet engine, development of a hypersonic ramjet engine is planned. To investigate performance of the intake of the hypersonic ramjet engine, wind tunnel test is done with freestream Mach number of 5.1. The total pressure recovery was 18 % with 12.9 % of ramp bleed. Several reasons for low total pressure recovery are shown. Supersonic internal compression is not enough. Then, the throat Mach number is high (M2.61) and total pressure losses at the terminal shock is large. Supersonic flow at the throat and position of the terminal shock is sensitive to a difference of the second ramp's throat height and the third ramp's throat height. Flow separations at the second ramp's trailing edge and the third ramp's leading edge are seen those could result in the trigger of unstart. The seal mechanism between the ramps and the sidewalls is important.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2010년도 제35회 추계학술대회논문집
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• pp.463-466
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• 2010

축열식 가열기 형태의 초음속 풍동을 이용하여 스크램제트 엔진 흡입구의 성능 연구를 수행하였다. 본 시험모델은 측벽배치변화가 성능에 큰 영향을 미치지 않는 것으로 나타났다. 격리부 성능연구에서는 Oblique shock train과 Normal shock train의 압력분포를 확인할 수 있었으며, 격리부의 불시동 한계성능을 분석하였다.

• Advances in aircraft and spacecraft science
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• 제5권3호
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• pp.349-362
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• 2018

A non-linear numerical simulation technique for predicting the unsteady performances of an airbreathing engine is developed. The study focuses on the simulation of integrated propulsion systems, where a closer coupling is needed between the airframe and the engine dynamics. In fact, the solution of the fully unsteady flow governing equations, rather than a lumped volume gas dynamics discretization, is essential for modeling the coupling between aero-servoelastic modes and engine dynamics in highly integrated propulsion systems. This consideration holds for any propulsion system when a full separation between the fluid dynamic time-scale and engine transient cannot be appreciated, as in the case of flow instabilities (e.g., rotating stall, surge, inlet unstart), or in case of sudden external perturbations (e.g., gas ingestion). Simulations of the coupling between external and internal flow are performed. The flow around the nacelle and inside the engine ducts (i.e., air intakes, nozzles) is solved by CFD computations, whereas the flow evolution through compressor and turbine bladings is simulated by actuator disks. Shaft work balance and rotor dynamics are deduced from the estimated torque on each turbine/compressor blade row.

• 대한기계학회논문집B
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• 제31권2호
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• pp.109-115
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• 2007

Airframe-integrated Scramjet engines of NASA Langley type consist of a compressor, a combustion chamber and a nozzle. When some disturbances occur in one module of the engine, its influences are propagated to other modules. In this study, it is investigated numerically how shock waves were caused by thermal choking in one module propagate upstream and how they influence adjacent modules. The calculations are carried out in 2-dimensional supersonic viscous flow model using explicit TVD scheme in generalized coordinates. The adverse pressure gradient caused by heat addition brings about separation of the wall boundary layers and formation of the oblique shock wave that proceed to upstream. This moving shock wave formed one module blocks the flow coming into the adjacent modules, which makes the modules unstarted.