本研究以研發真空推力50 lbf液體火箭引擎為目的，使用H2O2/W2JP推進劑組合，設定推進劑總質量流率為75 g/s、混合比(O/F)為4.0作為標準操作條件，採用三個液旋式噴注器組成噴注盤並設計燃燒室及噴嘴，以完成液態火箭引擎本體設計。為探討所設計之引擎性能，本研究共完成71次的大氣靜推力實驗，推進劑總質量流率範圍為61.3 g/s至104.5 g/s、混合比(O/F)範圍為2.1至6.1，其產生之燃燒室艙壓範圍為73.9 psia至257.0 psia、特徵速度範圍為1124 m/s至1831 m/s、實際量測到大氣推力範圍為21.1 lbf至42.1 lbf、比衝值範圍為131 s至216 s。

為了分析不同操作條件下的引擎性能，本研究透過統計方法，利用46筆實驗數據建立多元迴歸模型。模型設定推進劑總質量流率與混合比(O/F)作為操作變數，以預測引擎的比衝、推力、特徵速度及燃燒室艙壓。經由迴歸模型的顯著性檢定(F test)，結果顯示模型的信心水準皆大於99.9%，可在實驗操作範圍內用於預測引擎性能。

應用迴歸模型進行預測得到引擎在設計點的操作條件下(ṁ_Total：75 g/s；O/F：4.0)，大氣推力為31.4 lbf其對應比衝值為193 s。經由理想噴嘴計算得到，在噴嘴漸擴比為100並操作於真空環境中時，真空推力與真空比衝值分別為52.8 lbf與319 s，兩者皆分別高於原先設定值50.0 lbf與290 s。將不同推進劑總質量流率代入迴歸模型中進行預測，結果顯示在混合比(O/F)固定的情況下，隨著推進劑總質量流率提高，導致引擎推力與燃燒室艙壓上升，特徵速度與比衝則下降，合乎學理分析。

This study aimed to develop a liquid rocket engine with a vacuum thrust of 50 lbf, using an H₂O₂/W2JP propellant combination. The standard operating conditions were set with a total propellant mass flow rate of 75 g/s and a mixing ratio (O/F) of 4.0. The engine design incorporated a combustion chamber, nozzle, and an injector plate consisting of three liquid cyclonic injectors. To evaluate the performance of the designed engine, a total of 71 atmospheric static thrust tests were conducted. The total propellant mass flow rate ranged from 61.3 g/s to 104.5 g/s, with an O/F range of 2.1 to 6.1. The resulting combustion chamber pressures ranged from 73.9 psia to 257.0 psia, characteristic velocities from 1124 m/s to 1831 m/s, atmospheric thrust from 21.1 lbf to 42.1 lbf, and specific impulse values from 131 s to 216 s.

To analyze the engine's performance under various operating conditions, 46 experimental data points were used to establish a multiple regression model through statistical methods. The model utilized the total propellant mass flow rate and O/F ratio as operating variables to predict engine performance metrics, including specific impulse, thrust, characteristic velocity, and combustion chamber pressure. Based on the significance test (F-test), the regression model achieved a confidence level exceeding 99.9%, indicating its reliability for predicting engine performance within the experimental operating range.

Using the regression model to predict engine performance under the design point operating conditions (ṁ_Total：75 g/s；O/F：4.0), the atmospheric thrust was found to be 31.4 lbf, with a corresponding specific impulse of 193 s. Additionally, ideal nozzle calculations revealed that with a nozzle area expansion ratio of 100 operating in a vacuum environment, the vacuum thrust and specific impulse were 52.8 lbf and 319 s, respectively—both exceeding the initial design targets of 50.0 lbf and 290 s.

By substituting different total propellant mass flow rates into the regression model, predictions showed that, under a fixed O/F ratio, increasing the total propellant mass flow rate led to higher engine thrust and combustion chamber pressure, while the characteristic velocity and specific impulse decreased. These trends align with theoretical analyses.

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