隨著成本分散與風險分散的需求日增，各式太空飛行器的微小化成為近十年來相當熱門的研究方向。傳統衛星上所掛載的各系統元件之微小化，決定著如微衛星或皮米衛星等系統發展的成敗，而其中，在傳統衛星上扮演著舉足輕重地位的推進系統，則更是影響著微型化衛星系統之應用性的關鍵。
由於新製造技術的應用影響著衛星系統元件微小化的進程，本研究嘗試以微機電製造技術將傳統聯胺推進器進行微小化。考量下一個世代微衛星系統之需求，將所研發的微推進器之推力設定在千分之一牛頓等級，經由理論的估算與比較且評估微機電技術的材料相容性後，將微推進器的操作壓力與溫度設定在100psi及900K左右，而其對應的流量約為28.62mg/s，且在此條件限制進行聯胺微推進器之設計、製作與整合。
所設計之微推進器（尺寸約為8.48mm×4.0mm）包含了三個結構：隔離層、微反應器及微噴嘴。隔離層減少了約72%的接觸面積，以求減少由微反應器傳導至上游元件（如微控制閥）的熱傳量，除保護上游元件外，並期能降低微反應器之熱散失。在微反應器的設計上，為能有更大的彈性來處理不同流量以產生不同條件之熱氣源，重新設計的觸媒床為可增減之多層式結構，在本研究中，為能使得聯胺/觸媒反應得更完全來獲得較低的反應溫度，將實驗測試不同層數的觸媒床之氨分解率特性，並選擇適當之微反應器結構來整合成微推進器做最終之性能測試。在微噴嘴之設計上，考量可能的摩擦損耗將造成預期的推力無法達成，其喉部截面在實際設計製作之後，尺寸加大為60μm×77μm，而其擴張比則約為12。
在微反應器的測試上，利用氣相層析儀量得氨分解率的結果顯示，氨分解率隨著流量增加而降低的趨勢並不明顯，此外，觸媒床層數的增加對於氨分解率的影響亦不大，這除了是由於來自於氣體採樣上的誤差外，更多的是因為不論任何層數的觸媒床都提供了相當高的氨分解率（＞90%），使得趨勢顯得不明顯。基於這些測試結果，單層觸媒床的微反應器將被採用在後續的微推進器性能測試上。
微推進器性能的實驗測試結果顯示，在所測試的流量範圍內（0.34-0.91mg/s），所量得的推力位於0.43-1.33mN之間，實驗所得之推力與腔體壓力皆隨著流量而增加，然而，比衝值則在流量等於0.63mg/s處存在著一最大值（其推力約為1.01mN）。造成此現的原因推測為熱散失所造成之影響，流量增加造成氨分解率下降並使得反應溫度上升，從而造成更多的熱散失。
所建造之微推力測試台因結構上的影響，使得量測的推力在達平衡前有一近3秒之延遲，同一條件下腔體壓力之延遲僅有0.2秒，故該測試台僅能做一連續推力之量測，精確的脈衝型推力量測則需要對測試台做進一步的改良，且開發一與微推進器整合在一起的微控制閥來對推進劑的供給作更精準的控制。

MEMS techniques were adopted to miniaturize the conventional hydrazine monothruster in this thesis research. Based on the reaction control requirements of the micro- and nano-satellites, the targeted hydrazine microthruster was designed to provide millinewton-level thrusts. Restricted by MEMS materials and fabrication processes, the microthruster’s operation pressure and reaction temperature were set to be 100psi and 900K, respectively.

The designed microthruster was composed by a microreactor and a micronozzle, and was 8.48mm×4.0mm×2.0mm in size and weighted 0.14g. In order to reduce the heat transfer to the upstream device, an insulation layer was bonded on top of the microreactor with 72% contact area reduction. For the microreactor, the iridium-coated catalytic bed was designed and fabricated with a flow volume of 1.53mm3 and a shape of a high contact surface-to-volume ratio (84mm2/mm3), which provided an ammonia dissociation of 0.92 at the exhaust with an inlet hydrazine flow rate of 0.47mg/s. Considering the serious viscous losses in micro channel flow, the 3-D C-D micronozzle was designed to have a throat with a cross section of 60µm×77µm, and, the convergent and divergent area ratios were 3.3:1 and 12:1, respectively.
In the vacuum thrust tests, the design thruster showed thrusts of 0.43mN to1.33mN at the propellant flow rates from 0.34mg/s to 0.91mg/s. The best specific impulse of 162s occurred at the hydrazine flow rate of 0.63mg/s and the thrust of 1.01mN. Comparing to performance of conventional hydrazine thrusters, the analysis showed that the major degradation of the specific impulses of the designed microthruster were from the heat losses of the reactor.

The analyses of the onset of the chamber pressures showed a start-up delay time of about 134ms which was 2 to 5 folds of that of the macro scale hydrazine thrusters. It was believed that the requirement of a long induction period for the micro channel flow to reach it fully-developed condition account for the comparatively slow start-up trend of the microthruster.

This thesis research uniquely demonstrated the feasibility of developing a millinewton-level MEMS-based hydrazine thruster for space applications. However, due to the lack of appropriate microsensors, it was difficult to identify the detailed loss mechanism. Further studies are necessary to provide more detailed information to improve the design for the future development of the designed microthruster.

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