• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.327-330
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• 2006

Numerical simulations were carried out to investigate the base drag characteristics of a base bleed projectile with a central propulsive jet by considering the base homing process. Overall fluid dynamic process is modeled by Wavier-Stokes equations for reacting flows with two-equation $k-\omega$ SST turbulence closure. The combustion process is modeled by finite-rate chemistry with a given partially burned exit condition of the BBU (base-bleed unit). Besides the demonstrating the capability of the present CFD solver for the base drag and the interaction of the base flow with a rocket plume, present study gives an insight into the fluid dynamics and the combustion process of the hybrid-propulsion projectile.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.5
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• pp.445-455
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• 2010

In order to effectively predict thermo-acoustic instability within real combustors of rocket engines and gas turbines, in the present study, the Helmholtz equation in conjunction with the time lag hypothesis is discretized by the finite element method on three-dimensional hybrid unstructured mesh. Numerical nonlinearity caused by the combustion response term is linearized by an iterative method, and the large-scale eigenvalue problem is solved by the Arnoldi method available in the ARPACK. As a consequence, the final solution of complex valued eigenfrequency and acoustic pressure field can be interpreted as resonant frequency, growth rate, and modal shape for acoustic modes of interest. The predictive capabilities of the present method have been validated against two academic problems with complex impedance boundary and premixed flame, as well as an ambient acoustic test for liquid rocket combustion chamber with/without baffle.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.8
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• pp.699-705
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• 2007

Recent experimental data shows that the noticeable feature of irregular roughened spots on the fuel surface occurs during the combustion test with PMMA/GOX in the hybrid rocket motor. The generation of these unexpected patterns is likely to be resulted from the disturbed boundary layer due caused by wall blowing which is intented to simulate the process of fuel vaporization. LES technique was implemented to investigate both the flow characteristics near fuel surface and the subsequent evolution of turbulence modified by the wall blowing. Simple channel geometry instead of circular grain configuration was used for the investigation without chemical reactions in order to allow for a focused examination on the near-wall behavior at the Reynolds number of 22,500. It was shown that the wall blowing pushed turbulent structures upwards making them tilted and this skewed displacement, in effect, left the foot prints of the structures on the surface. This change of kinematics may explain the formation of irregular isolated spots on the fuel surface observed in the experiment.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.210-213
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• 2010

Recent experimental data shows that the noticeable feature of irregular roughened spots on the fuel surface occurs during the combustion test. The generation of these unexpected patterns is likely to be resulted from the disturbed boundary layer due caused by wall blowing which is intended to simulate the process of fuel vaporization. LES without chemical reaction was conducted to investigate the flow characteristics at the near-fuel surface and the behavior of turbulent structures which is evolved by the wall blowing at the Reynolds number of 23,000. Cylindrical geometry was considered to get the most reality of the calculation results because real hybrid rocket motor is circular grain configuration. It was shown that the wall blowing pushed turbulent structures upwards making them tilted and this skewed displacement, in effect, left the foot prints of the structures on the surface. This change of kinematics may explain the formation of irregular isolated spots on the fuel surface observed in the experiment.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.7
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• pp.585-592
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• 2016

Visualization was done by using $CH^*$ chemiluminescence images and PMT measurements in order to understand the origin of fluctuating pressure and chemical luminosity at about 500 Hz frequency even in stable combustion, which was observed in recent experimental tests, and to find the physical correlation leading to Low Frequency Instability(LFI) in terms of phase angle. In stable combustion, chemical reactions are distributed along the shear layer flow showing a negative coupling(about 180 degree in phase angle) with combustion pressure. However, phase difference is shifted to a positive coupling showing less than 90 degree in unstable case. Also a periodic change in the distribution of chemical reactions is observed along with local flame extinction and the appearance of big scale vortex flow. In the transition to LFI, local flame extinction and small vortex flow start to appear in a row. As seen in the bluff body wake in reactive flow, the periodic appearance of vortex flow seems to share the same physical process of BVK(Bernard Von Karman) instability generation. Thus, the appearance of local extinction in 500 Hz fluctuations is gradually amplified to complete extinctions of about 20 Hz, and it leads into LFI.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.243-247
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• 2007

Fuel mass flux was experienced with a function of the oxidizer mass flux using initial port area of solid fuel, in stead of regression rate correlation which shows combustion characteristic in hybrid propulsion. The burning rate could be easily obtained by using the oxidizer mass flux of initial port area without iteration, and fuel configuration could be designed simply. In this experiments PE was used as fuel, COX was used as oxidizer. A variation of mass flux of solid fuel with port area is considered by changing the burning time. In the case of approximate 0.5 for an exponent of oxidizer mass flux, using the fuel mass flux correlation is more suitable than regression rate correlation in hybrid propulsion.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.368-372
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• 2005

The variation of regression rate of single port hybrid rocket was studied with various fuel. As fuel, PE, PMMA were used and gas oxygen as oxidizer. The regression rate depends on but flow rates of oxidizer also thermodynamic properties of fuel. In this study, the empirical relation for regression rate of solid fuel were found with mass transfer coefficient(B number) and oxidizer flow rate.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.1
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• pp.64-72
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• 2010

Due to the interaction between main oxidizer flow and the wall injected flow resulting from the regression process, a specific time characteristics identified in the frequency spectrum of streamwise velocity is generated in the hybrid rocket motor. In order to understand the response of the turbulent flow to two different types of external momentum forcing, LES analysis was conducted without considering the combustion. It turns out that both concentrated and distributed forcings do not lead to the disastrous resonance phenomenon. Energy contents are enhanced due to the added momentum but the peak frequency was not modified in the turbulent flow near the end of the rocket motor. Natural frequency of the flow system should be taken into account to further pursue the instability issue by using external forcing.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.1
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• pp.62-68
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• 2009

The interaction between wall blowing and oxidizer flow can generate a very complicated flow characteristics in combustion chamber of hybrid rockets. LES analysis was conducted with an in-house CFD code to investigate the features of turbulent flow without chemical reactions. The numerical results reveal that the flow oscillations at a certain frequency exists on the fuel surface, which is analogous to those observed in the solid propellant combustion. However, the observation of oscillating flow at a certain frequency is only limited to a very thin layer adjacent to wall surface and the strength of the oscillation is not strong enough to induce the drastic change in temperature gradient on the surface. The visualization of fluctuating pressure components shows the periodic appearance of relatively high and low pressure regions along the axial direction. This subsequently results in the oscillation of flow at a certain fixed frequency. This implies that the resonance phenomenon would be possible if the external disturbances such as acoustic excitation could be imposed to the oscillating flow in the combustion chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.244-245
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• 2003

고체나 액체 추진로켓에 비하여 하이브리드 추진 시스템은 작동조건의 안정성과 안전함등의 많은 장점을 가지고 있다. HTPB와 같은 고체연료는 제작 및 저장, 운송 그리고 장착상의 안정성을 가지고 있으며 하이브리드 로켓의 고체연료로의 산화제의 유입을 제어하면서 추력의 변화와 엔진내부의 연소중단과 재 점화를 용이하게 할 수 있다. 이러한 이유로 인하여 하이브리드 엔진은 좀 더 경제적인 장치로 기대를 모으고 있다. 그러나, 기존의 하이브리드 로켓 엔진은 고체 추진 로켓에 비하여 낮은 연료 regression 율과 연소효율을 가지는 단점이 있다. 이러한 단점을 해결하고 요구되어지는 추력값과 연료유량을 증가시키기 위하여 고체연료의 표면적을 증가시킬 필요가 있다. 기존의 하이브리드 엔진에서는 연료 그레인에 다수의 연소포트를 만들어 표면적을 증가시켰으나 이는 비 활용 공간의 증가와 추진제의 질량 및 체적분율의 상당한 감소를 초래한다. 지난 수십년간에 걸쳐 하이브리드 엔진에서 연료의 regression 특성 및 엔진 성능 향상을 위한 연구가 계속되어 왔으며 최근에 엔진의 체적 규제를 경감시키고 연료의 regression율을 향상시키기 위하여 선회유동을 이용하는 하이브리드 로켓 엔진들이 제안되고 있다. 이러한 선회유동을 가지는 하이브리드 로켓은 고체연료 그레인에 대하여 평행하게 유입되는 기존의 하이브리드 로켓에 비하여 고체연료 벽면에서의 대류열전달이 현저하게 증가하게 되어 아주 높은 고체연료의 regression율을 얻을 수 있는 이점이 있다. 선회유동 하이브리드 로켓의 연소과정은 고체 연료의 열분해과정, 대류 열전달, 난류 혼합, 난류와 화학반응의 상호작용, soot의 생성 및 산화과정, soot 입자 및 연소가스에 의한 복사 열전달, 연소장과 음향장의 상호작용 등의 복잡한 물리적 과정을 포함하고 있다. 이러한 물리적 과정 중 난류연소, 고체연료 벽면 근방에서의 대류 열전달 및 연소과정에서 생성되는 soot 입자로부터의 복사 열전달, 그리고 고체연료 열 분해시 표면반응들은 고체연료의 regression율에 큰 영향을 미친다. 특히 고체연료의 난류화염면의 위치와 폭, 그리고 비 예혼합 난류화염장에서 생성되는 soot의 체적분율의 예측은 난류연소모델, 열전달 모델, 그리고 regression율 모델에 의해 크게 영향을 받기 때문에 수치모델의 예측 능력 향상시키기 위하여 이러한 물리적 과정을 정확히 모델링해야 할 필요가 있다. 특히 vortex hybrid rocket내의 난류연소과정은 아래와 같은 Laminar Flamelet Model에 의해 모델링 하였다. 상세 화학반응 과정을 고려한 혼합분율 공간에서의 화염편의 화학종 및 에너지 보존 방정식은 다음과 같다. 화염편 방정식과 혼합분률과 scalar dissipation rate의 관계식을 이용하여 혼합분률과 scalar dissipation rate에 따른 모든 reactive scalar들을 구하게 된다. 이러한 화염편 방정식들을 mixture fraction space에서 이산화시켜서 얻은 비선형 대수방정식은 TWOPNT(Grcar, 1992)로 계산돼 flamelet Library에 저장되게 된다. 저장된 laminar flamelet library를 이용하여 난류화염장의 열역학 상태량 평균치는 presumed PDF approach에 의해 구해진다. 본 연구에서는 강한 선회유동을 가지는 Hybrid Rocket 연소장내의 난류와 화학반응의 상호작용을 분석하기 위하여 Laminar Flamelet Model, 화학평형모델, 그리고 Eddy Dissipation Model을 이용한 수치해석결과를 체계적으로 비교하였다. 또한 Laminar Flamelet Model과 state-of-art 물리모델들을 이용하여 선회 유동을 갖는 하이브리드 로켓 엔진의 연소 및 Soot 생성 및 산화과정을 살펴보았으며 복사 열전달이 고체 연료 표면의 regression율에 미치는 영향도 살펴보았다. 특히 swirl강도, 산화제의 유입위치 그리고 선회유동의 형성방식이 하이브리드 로켓의 연소특성 및 regression rate에 미치는 영향을 상세히 해석하였다.