本論文配合國家太空中心(NSPO)探空火箭六號「衛星推進控制系統(RCS)酬載」之開發，自行設計、測試並性能分析下吹式(blow-down type)聯胺推進系統。系統主要以高壓氮氣儲瓶、聯胺燃料槽(propellant tank)、止回閥(check valve)、爆通閥、過濾器(filter)、工作閥(service valve)、推進器控制閥(thruster valve)及聯胺推進器(thruster)等元件組成，其中自行設計之聯胺單基推進器採單管徑向式噴注器設計，並使用兩段式觸媒床填充方式（14~18mesh、20~25mesh），應用檔板及彈簧以分隔及保持觸媒充填密度。設計錐形推力噴嘴擴張比為40:1，預期之真空推力約為1lbf。
　　經真空靜推力測試顯示，控制聯胺流率為1.97g/sec時，連續操作模式之反應室壓力達202psi，產生1.06lbf真空推力。經計算得比衝值約243sec、特徵速度為4090ft/sec、噴嘴效率(CF)為1.91。在250ms脈衝推力測試顯示此推進系統之平均衝量為0.28±0.01lbf．sec；在500ms脈衝測試之平均衝量為0.51±0.01lbf．sec，皆具穩定之衝量輸出，符合衛星姿態控制需求。
　　本研究之系統性能測試顯示反應室及管路壓力震盪，經實驗研究分析顯示：
(1)本研究設計之聯胺推進系統反應室及管路壓力震盪起源於推進器內之劇烈化學反應。
(2)觸媒床擋板彈簧對於反應室及管路壓力震盪有明顯的影響；選用較小彈性係數（K值）之彈簧，對於吸收反應室之壓力震盪具較佳之功效。
(3)管路內配置0.5mm口徑之孔口及空腔具延遲壓力震盪現象(4-5sec)之功能，但乃未能完全抑制壓力震盪。
　　本研究未來將針對推進器內之觸媒填充方式及噴注器設計進行探討，期改變反應室內之化學反應模式及壓力震盪特性；並探討配置孔口及空腔所產生之延遲壓力震盪原因，以確立推進系統穩定壓力輸出。

In conjunction with the payload research project of the Sounding Rocket VI mission of NSPO, this thesis designed an 1-lbf propulsion system and tested in atmosphere and in vacuum environment. This blow-down type propulsion system composed of a pressurant’s tank, a propellant tank, check valves, an explosive valve, filters, service valves, and a pair of hydrazine thrusters. The self-designed hydrazine thruster was equipped with a radial-type injector. Two different sizes (14~18mesh、20~25mesh) Shell 405 catalyst were packed in the reaction bed. The catalysts were separated and held by porous plates against a spring to keep the catalyst density constant. Conical nozzle with 40:1 expansion ratio was used to produce a thrust output near 1lbf.
At the static thrust tests in a vacuum chamber, the system produced a chamber pressure of 202psi and a thrust of 1.06lbf in continue hydrazine flow rate of 1.97g/sec. The specific impulse　(Isp) and characteristic velocity (C*) were calculated to be 243sec and 4090ft/sec, respectively. In the pulse operations of 250ms and 500ms durations, the average impulses were 0.28±0.01lbf-sec and 0.51±0.01lbf-sec. The performance of the propulsion system was justified to meet the requirements for a reaction control system of satellites.
The results of the test showed pressure oscillation in the reactor and thrust oscillation as well during hot-fire operations. The phenomenon may induce inaccuracy while applying the system to the control of satellites. Further investigation of the this instability (pressure oscillation) showed
(1)the source of the oscillations was from the intensive chemical reaction processes;
(2)a spring with a relatively small spring constant showed effect on vibration absorption that suppressed the pressure oscillation in the reactor;
(3)a restricting orifice of o.5mm ID mounted in the upstream of the thruster valve showed significant effects on the delaying (4-5 seconds) and suppressing the pressure oscillations.
For future works, the research will focus on the effect of catalyst packing, the design of the injector, usages of cavity and orifices on suppression of the pressure oscillation to ensure a stable performance of this hydrazine propulsion system.

中文摘要 I
英文摘要 III
誌謝 V
目錄 VI
表目錄 VII
圖目錄 VIII
符號說明 X
第一章　緒論 1
1-1 前言 1
1-2 研究動機及目的 1
1-3 文獻回顧 2
第二章　反應特性與推進器性能 10
2-1 燃料與觸媒特性 10
2-2 推進器性能 13
第三章　推進系統設計與測試 15
3-1 聯胺推進器之設計 15
3-2 推進系統設計與測試 19
第四章　系統不穩定分析 26
4-1 實驗方法 26
4-2 實驗設備 28
4-3 實驗結果與分析 29
4-4 討論 34
第五章　緒論與未來工作 36
5-1 結論 36
5-2 建議及未來研究方向 37
參考文獻 38
附錄 92
自述 103

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