• 대한기계학회논문집B
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• 제38권12호
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• pp.1009-1015
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• 2014

본 연구는 NACA0012 천음속 에어포일 유동에 있어서 비평형 응축이 Force 계수(압력, 양력 및 항력계수)에 미치는 영향을 TVD 수치해석을 통하여 연구하였다. 정체점 온도 298 K, 받음각 ${\alpha}=3^{\circ}$인 경우, 주류 마하수 0.78~0.81에서는 정체점 상대습도의 증가함에 따라 양력은 단순 감소한다. 반면 Lift force break 마하수 영역의 주류 마하수에서는 정체점 상대습도의 증가에 따라 양력은 오히려 증가한다. 받음 각 ${\alpha}=3^{\circ}$, 정체점 상대습도가 0%인 경우, 주류 마하수의 증가에 따라 항력은 급격하게 증가하지만, 응축의 영향이 큰 60%인 경우에는 주류 마하수의 증가에 조금 증가할 뿐이다. 동일한 주류 마하수인 경우 비평형 응축에 따른 전 항력의 감소는 받음각과 정체점 상대습도가 증가할수록 크게 된다. 응축이 없는 ${\Phi}_0=0%$인 경우는 주류 마하수가 크고 받음각이 클수록 Wave drag은 크게 되나 응축의 영향이 비교적 큰 ${\Phi}_0=50%$ 이상인 경우는 오히려 Wave drag이 작아지는 것으로 나타났다. 한편, 정체점 상대습도가 낮고, 주류 마하수가 클수록 충격파 직전의 최대 마하수는 커지는 것으로 나타났다.

• 한국항공우주학회지
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• 제48권5호
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• pp.323-334
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• 2020

경계층 천이는 극초음속 비행체의 열공력 성능에 직접적인 영향을 미치기 때문에, 성능 해석과 설계를 위해 천이지점의 정확한 예측 능력이 필수적으로 요구된다. 본 연구에서는 극초음속 경계층 천이 예측을 위한 γ-Re_θt 모델을 개선하고 검증하였다. 천이 시작 위치 결정에 영향을 주는 간헐도 수송방정식 생성항의 계수를 압축성 경계층 상사해를 기반으로 마하수와 벽면온도, 자유류 정체온도 등에 대한 함수 형태로 구성하고 적용하였다. 기본 모델의 Re_θc와 F_length 상관관계식에 마하수에 따른 천이 시작 운동량두께 레이놀즈수와 천이구간의 길이 변화를 반영할 수 있도록 마하수에 대한 관계식을 추가로 결정하여 적용하였다. 실용적인 사용을 고려하여 제안 모델을 상용 CFD 코드에 적용하였으며, 검증을 위해 모델을 사용하여 극초음속 평판과 원뿔 경계층 해석을 수행하였다. 실험 결과와의 비교를 통해 마하수와 단위 레이놀즈수 변화에 대한 개선된 예측성능을 확인하였다.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2003년도 제20회 춘계학술대회 논문집
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• pp.91-93
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• 2003

A comprehensive numerical study is carried out to investigate for the understanding of the flow evolution and flame development in a supersonic combustor with normal injection of ncumally injecting hydrogen in airsupersonic flows. The formulation treats the complete conservation equations of mass, momentum, energy, and species concentration for a multi-component chemically reacting system. For the numerical simulation of supersonic combustion, multi-species Navier-Stokes equations and detailed chemistry of H2-Air is considered. It also accommodates a finite-rate chemical kinetics mechanism of hydrogen-air combustion GRI-Mech. 2.11[1], which consists of nine species and twenty-five reaction steps. Turbulence closure is achieved by means of a k-two-equation model (2). The governing equations are spatially discretized using a finite-volume approach, and temporally integrated by means of a second-order accurate implicit scheme (3-5).The supersonic combustor consists of a flat channel of 10 cm height and a fuel-injection slit of 0.1 cm width located at 10 cm downstream of the inlet. A cavity of 5 cm height and 20 cm width is installed at 15 cm downstream of the injection slit. A total of 936160 grids are used for the main-combustor flow passage, and 159161 grids for the cavity. The grids are clustered in the flow direction near the fuel injector and cavity, as well as in the vertical direction near the bottom wall. The no-slip and adiabatic conditions are assumed throughout the entire wall boundary. As a specific example, the inflow Mach number is assumed to be 3, and the temperature and pressure are 600 K and 0.1 MPa, respectively. Gaseous hydrogen at a temperature of 151.5 K is injected normal to the wall from a choked injector.A series of calculations were carried out by varying the fuel injection pressure from 0.5 to 1.5MPa. This amounts to changing the fuel mass flow rate or the overall equivalence ratio for different operating regimes. Figure 1 shows the instantaneous temperature fields in the supersonic combustor at four different conditions. The dark blue region represents the hot burned gases. At the fuel injection pressure of 0.5 MPa, the flame is stably anchored, but the flow field exhibits a high-amplitude oscillation. At the fuel injection pressure of 1.0 MPa, the Mach reflection occurs ahead of the injector. The interaction between the incoming air and the injection flow becomes much more complex, and the fuel/air mixing is strongly enhanced. The Mach reflection oscillates and results in a strong fluctuation in the combustor wall pressure. At the fuel injection pressure of 1.5MPa, the flow inside the combustor becomes nearly choked and the Mach reflection is displaced forward. The leading shock wave moves slowly toward the inlet, and eventually causes the combustor-upstart due to the thermal choking. The cavity appears to play a secondary role in driving the flow unsteadiness, in spite of its influence on the fuel/air mixing and flame evolution. Further investigation is necessary on this issue. The present study features detailed resolution of the flow and flame dynamics in the combustor, which was not typically available in most of the previous works. In particular, the oscillatory flow characteristics are captured at a scale sufficient to identify the underlying physical mechanisms. Much of the flow unsteadiness is not related to the cavity, but rather to the intrinsic unsteadiness in the flowfield, as also shown experimentally by Ben-Yakar et al. [6], The interactions between the unsteady flow and flame evolution may cause a large excursion of flow oscillation. The work appears to be the first of its kind in the numerical study of combustion oscillations in a supersonic combustor, although a similar phenomenon was previously reported experimentally. A more comprehensive discussion will be given in the final paper presented at the colloquium.