混合火箭結合了固態火箭以及液態火箭的優點，具有安全性高、可靠度佳、價格便宜、推力之可控制性以及對環境影響較低等特性，然而因傳統混合火箭較低燃料退縮率致使其性能不易提升，限制其使用範圍。本研究利用以石蠟混合HTPB方式以改進石蠟燃料受熱後強度問題，並以漩渦進氧以提高氧化劑/燃料蒸氣混合機制，探討石蠟基燃料退縮率及燃燒效率等特性。實驗中依重量百分製作100P(純石蠟)、90P(90%石蠟+10% HTPB)、70P及50P等燃料，高石蠟含量燃料如90P受熱後藥柱強度即明顯不足，70P燃料則在較高氧氣通量時，其燃燒室壓力出現振盪，且有部份熔融燃料隨噴焰噴出。至於50P燃料藥柱受熱後仍可維持一定強度，對石墨噴嘴燒蝕現象較為趨緩，其燃料退縮率在漩渦進氣條件下可表示r=0.026GOx^0.8076，在GOx=150Kg/m^2-sec時，其燃料退縮率較傳統混合火箭HTPB/GO2 增加61% ，此種50P之石蠟基燃料可同時滿足提升燃料退縮率及降低噴嘴燒蝕等研究目標。檢測試驗後藥柱燃面，發現在燃料蒸氣與氧氣混合機制較差之燃燒情況下，藥柱通道出口處燃料退縮率急遽上升，顯示部份未能於燃燒室反應之混合燃氣繼續於後混室燃燒。因此，對於混合機制較為不足系統後混室為必需的裝置。研究中亦根據燃燒室壓力及推力量測數據及燃料之特徵排放速度C*，並利用質量守恆原理，推估混合火箭燃燒過程中燃料退縮率、氧化劑通量、比衝值、氧化劑/燃料比 (O/F)、推力係數等內彈道性能變化，藉以了解混合火箭複雜燃燒行為，進而設計高性能混合火箭滿足下世代任務需求。經質量守恆之內彈道性能分析，可於一次實驗數據中獲得大量燃料退縮率與氧氣通量關係，此關係式尤適用在與實驗平均氧氣通量接近範圍。此質量守恆分析方法，可提供作為預估燃料退縮率之有效方法。

The hybrid rocket has the superiority of safety, low-cost and high-reliability over solid-propellant and liquid-propellant rockets. Classical hybrids, however have suffered from slow fuel regression rates and relatively poor combustion efficiency. Effects of fuel composition and swirling oxidizer injection on the fuel regression rates were investigated. Various weight ratios of paraffin wax and HTPB mixtures including 90P (90% paraffin + 10% HTPB), 70P and 50P were developed. Grain strength was dropped significantly as test motor was ignited for the fuel of 90P. 70P fuel could be maintained stable combustions under most test conditions except at higher oxygen flux test in which pressure fluctuations was observed, and part of fuel vapor was blow with exhaust flame. With swirling GOx injection, 50P fuel has shown high regression rate and grain strength properties.Variation of regression rate with gaseous oxygen could be expressed as r=0.026GOx^0.8076, which is about 61% increased as compared with classical HTPB/GO2 hybrid rocket at GOx=150Kg/m^2-sec . Moreover, 50P fuel motor could also be operated at fuel rich combustion condition that would reduce graphic nozzle erosion effect. The post test grain contour was employed to estimate regression rate distribution along axial direction. A rapid increased regression rate near port end was observed in lower mixing conditions that were caused by low oxygen mass flux. It indicated that fuel vapor and oxidizer mixture gas was unable to mix well and react completely inside the combustor and part of the combustible gas was therefore, burnt continuously in the aft-mixing chamber. Aft-mixing chamber was necessary for the system lacking of strong mixing mechanism. With the measured chamber pressure and oxygen mass flow rate and the calculated fuel characteristic velocity C*, histories of the spatially averaged of regression rate, oxygen mass flux, oxidizer/fuel ratio, and thrust coefficient were obtained by employing mass conservation law. Due to the variation of port diameter, wide range of oxygen mass flux and regression rate could be obtained during single test. This analysis has proven as an effective way to estimate fuel regression rate. Analysis on combustion history inside a motor will probably provide as a guideline for designing a high performance hybrid rocket.

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