本論文針對小型雙推進劑MMH/NTO液態火箭引擎燃燒室，以數值方法模擬其燃燒流場。並探討在總燃料(MMH)噴注量固定不變之情形下，使用部分燃料噴向燃燒室壁來冷卻燃燒室壁面，對燃燒室之燃燒流場及壁面冷卻的影響。其中探討之可變參數包括：冷卻燃料噴注量之比例、冷卻用燃料碰撞壁面後之擴散角度及冷卻燃料注油之角度。分析結果發現，在本文之注油器型態之下，30%之冷卻燃料噴注比例，造成與NTO衝擊霧化之MMH減少，導致整體液滴的混合不均及燃燒效率不佳。而10%之冷卻燃料噴注比例造成與NTO衝擊霧化之MMH增加，使得燃燒較為集中於流場中心，並提高燃燒效率。在冷卻燃料之擴散角度方面，增加擴散角度有助於加大冷卻擴散面包覆燃燒室的範圍，對於燃燒室壁前端冷卻有較為顯著之影響。另外，增加冷卻燃料之噴注角度，會使得冷卻燃料提早碰撞燃燒室壁，而提早產生冷卻擴散面，降低燃燒室壁前端之局部高溫。此外，衝擊霧化正面擴散角度之大小亦會對燃燒室壁溫度產生影響，相較於60度之擴散角， 90度之擴散角度會導致散佈在霧化液膜外圍之液滴直接噴撞燃燒室壁，造成燃燒室壁前端產生高溫的現象。

In this thesis, the combustion flow in a small bipropellant liquid rocket combustor is numerically simulated. Effects of injecting partial fuel on the chamber wall for cooling are investigated. The parameters investigated include the ratio of the cooling fuel to the total injected fuel, the spray angle of the cooling fuel and the injection angle of the cooling fuel. The results obtained from the simulation show that the 30% ratio of the cooling fuel reduces the flow rate of MMH impinging with NTO, resulting in uneven mixing and lower combustion efficiency. The 10% ratio of the cooling fuel makes the combustion area concentrate in the center region of the combustor, increasing the combustion efficiency. Increasing the spray angle of the cooling fuel leads to a larger cover area of the cooling fuel on the chamber wall. The fore-end of the chamber wall is thus cooled more effectively. Additionally, a larger injection angle reduces the peak temperature of the fore-end wall of the combustor. Compared to the cases with a 60o atomization spray angle, the cases with a 90o atomization spray angle cause more droplets in the outer region of the spray flowing towards the chamber wall, resulting in higher temperature in the fore-end wall of the chamber.

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