액체메탄-액체산소 소형 로켓 연소실의 연소특성에 관한 수치적 연구 = Numerical Study on the Combustion Characteristics of LCH4-LOx Small Rocket Combustor

다국어 초록 (Multilingual Abstract)

With the increasing demand for reusable and cost-effective space launch vehicles, liquid methane (LCH4) has emerged as a promising propellant owing to its high specific impulse, non-toxicity, in-situ resource utilization (ISRU) compatibility, and minimal coking characteristics suitable for reusable engines. However, the combustion of LCH4-LOx involves complex multi-physics phenomena, including cryogenic liquid injection, atomization, vaporization, turbulent mixing, and chemical reactions. Therefore, developing a reliable and efficient numerical methodology is essential for optimal engine design and performance prediction. In this study, three-dimensional computational fluid dynamics (CFD) simulations were conducted to investigate the combustion characteristics and flow structures of a small LCH4-LOx rocket combustor. The computational domain was extended from the injector recess through the combustion chamber and converging-diverging nozzle to the external plume region, enabling simultaneous observation of internal combustion phenomena and external supersonic flow characteristics including shock diamond structures. To accurately capture the turbulent reacting flow, the Reynolds-Averaged Navier-Stokes (RANS) equations were solved with the k-ω SST turbulence model, and the Eddy Dissipation Concept (EDC) model was employed for turbulence-chemistry interactions. A detailed chemical mechanism (DRM-19) consisting of 21 species and 84 reactions was implemented to resolve the methane-oxygen combustion process. Advanced acceleration techniques including Chemistry Agglomeration, Dynamic Mechanism Reduction (DMR), and In-Situ Adaptive Tabulation (ISAT) were introduced to mitigate the computational expense associated with detailed chemistry. The multiphase flow characteristics of cryogenic propellants were simulated using the Eulerian-Lagrangian approach, where the continuous gas phase was solved with the Eulerian framework and the discrete liquid droplets were tracked in a Lagrangian manner. The Wave breakup model was applied to simulate secondary atomization processes. As a results, the temperature field analysis revealed distinct swirl-induced combustion structures characterized by a low-temperature core along the centerline and high-temperature zones near the chamber walls. As the equivalence ratio increased, the high-temperature region extended upstream and distributed more continuously throughout the combustion chamber. Species distribution analysis showed that increasing equivalence ratio enhanced fuel-oxidizer mixing efficiency, resulting in improved characteristic velocity. At the lowest equivalence ratio of 1.20, significant amounts of unreacted oxygen were observed even at the nozzle exit, indicating incomplete mixing. In contrast, the highest equivalence ratio of 2.06 exhibited the widest and most intense flame zone, as indicated by OH radical distribution, effectively utilizing the combustor volume. Flow field analysis identified strong recirculation zones at the upstream corner regions, which contributed to flame stabilization and promoted droplet atomization and mixing. The strongest recirculation structure was observed at equivalence ratio 2.06, forming a clear concentric vortex pattern. In the external flow region, shock diamond patterns including Mach disks were clearly captured, demonstrating the capability of the numerical method to resolve complex supersonic expansion flows. This study presents a validated and efficient combustion analysis framework incorporating detailed chemical kinetics for small LCH4-LOx rocket engines. The developed methodology successfully captures the complex interactions among turbulent flow, spray dynamics, and finite-rate chemistry with reasonable computational cost. The findings provide valuable insights into the effects of equivalence ratio on mixing efficiency, combustion intensity, and thrust performance. The proposed numerical approach can serve as a practical design tool for injector geometry optimization, equivalence ratio selection, and thermal management in the development of next-generation methane-fueled propulsion systems.

목차 (Table of Contents)

• Ⅰ. 서 론 1
• 1. 연구 배경 및 목적 1
• 2. 연구 동향 5
• Ⅱ. 수치적 방법론 8
• 1. 물리적 모델 8
• 2. 난류 모델 10
• 3. 화학 반응 12
• 3.1 난류-연소 상호작용 모델 13
• 3.2 반응 메커니즘 15
• 3.3 화학 반응 계산을 위한 가속화 기법 20
• 4. 다상 유동 22
• 4.1 입자 운동 24
• 4.2 열 및 질량 교환 26
• 4.3 2차 분열(secondary breakup) 29
• 5. 이산화 31
• 6. 해석 알고리즘 34
• Ⅲ. 해석 조건 36
• 1. 계산영역 36
• 2. 경계 및 분사 조건 37
• 3. 격자계 40
• Ⅳ. 해석 결과 43
• 1. 연소 및 추력 성능지표 43
• 2. 연소실 내부 온도 분포 45
• 3. 연소실 내부 화학종 분포 47
• 3.1 산화제: 산소(O2) 47
• 3.2 연료: 메탄(CH4) 49
• 3.3 연소 생성물: 물(H2O), 이산화탄소(CO2), OH 라디칼 51
• 4. 연소실 및 초음속 노즐의 후류 영역 유동 57
• Ⅴ. 결 론 60