本論文研究完成一組20牛頓級NTO/MMH推進器之設計與製造，並以地面靜推力測試進行推進器之性能驗證。推進器可概分為噴注盤、燃燒室與噴嘴，本研究以噴注盤設計為主，分別完成二對雙衝擊噴流設計及單對三衝擊噴流設計。設計以冷流實驗為起步，在不同之環境壓力及噴注孔徑進行PLIF實驗，觀察推進劑模擬液衝擊霧化後之分布，以選定適當噴注孔徑組合。本研究並依據霧化分布特性及開放空間燃燒特性設計燃燒室，同時為避免壁溫過高，燃燒室採用液膜冷卻設計。推進器採用鐘型噴嘴(Bell shape Nozzle)，以理論基礎估算噴嘴幾何尺寸，並利用商業軟體(Nozzle3.7)計算噴嘴外型。
本研究完成三衝擊噴流噴注盤製造，整合燃燒室與噴嘴後進行地面靜推力實驗。當推進器操作於接近設計流量時(推進劑7.17g/s、混合比1.27)，實驗結果顯示燃燒室壓力約3.96bar、地面測量推力7.9N、比衝值約113s；推算之真空推力為19.3N、比衝值約275s，初步達成20牛頓級之設計目標。此外，燃燒室壁面溫度大多控制介於100°C ~ 400°C，顯示冷卻設計具有功效；根據實驗資料回歸分析顯示於1.4~1.6之操作混合比，推進器具有較佳之比衝值，當混合比為1.7時，推進器具有較佳之特徵速度，故未來可降低壁面冷卻流量，提高整體操作混合比，進而提升推進器性能(比衝值、特徵速度)。噴嘴效率方面，以地面量測推力及理想地面推力估算，顯示噴嘴效率約74.5%，其推力系數約1.24，未來噴嘴設計部分仍有改善之空間。

This thesis research has completed the design and fabrication and test of a 20N NTO/MMH thruster which was composed by an injector plate, combustor and nozzle. This study focused mainly on the injector plate design, a two-pair doublet impinging injector design and a triplet impinging injector design had been completed. PLIF technique was adopted in cold-flow experiments to observe the mass distributions of the simulants of propellants and the mixing characteristics of the impinging sprays from the injectors with different injector orifices and ambient pressures. Based on the observations, appropriate orifice size combinations were selected for the injector plate design. In conjunction with the previous hot fire observations, the dimensions of the combustors were also determined from the cold-flow observations. In addition, film cooling of the combustor wall was also designed. For the thrusters, bell-shape nozzle was applied, and the nozzle geometry was estimated by fundamental theory and a basic design software (Nozzle3.7).
This research has completed the fabrication of a thruster with a triplet impinging injector plate. Sea-level static thrust tests were executed with the thruster. Near the design point (mass flow rate of 7.17g/s and O/F of 1.27), the thruster produced a chamber pressure of 3.96bars and wall temperatures within 400K to 700K. The measured sea-level thrust was 7.9Newtons and a specific impulse of 113seconds was deduced. An ideal vacuum thrust and specific impulse were estimated to be 19.3Newtons and 275seconds with the measured chamber pressure and assumed chamber temperature (2000K). Experiments also showed that better specific impulses at O/Fs from 1.4 to 1.6 and a better characteristic velocity at O/F of 1.7 could be deduced, that is, a reduction of the film cooling MMH flow rate from the original design would raise the propulsive performance of the thruster. In addition, the nozzle efficiency and the thrust coefficient were estimated to be 74.5% and 1.24, respectively. This indicated that a more detailed and realistic design of the nozzle is crucial in order to improve the performance of the thruster.

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