臺灣博碩士論文加值系統

本研究專注於發展四軸混合式火箭引擎推進平台之推進系統。四軸混合式火箭引擎推進平台依其特性及任務需求劃分為以下三個次系統：推進、結構及航電，其中航電系統又包含導航與控制、通訊及供電系統。由於混合式火箭引擎擁有安全性、可存放性、成本低廉與取得容易及可控性等優勢，四軸火箭引擎推進平台選擇了混合式火箭引擎做為推進之核心。混合式火箭通常使用液態氧化劑及固態燃料做為推進劑，因此設備需求包含高壓氣瓶、管路及燃燒艙。本研究使用了氮氣（N2）做為加壓氣體、90 wt%高濃度雙氧水（H2O2）做為氧化劑以及固態聚丙烯做為燃料。
為了達成四軸混合式火箭引擎推進平台之規格需求，推進系統主要目標為提供240公斤全推力並且在全推力的狀態下達到30秒的續航力。本系統共搭載四顆引擎和六個10公升複合材料高壓氣瓶，其中兩組高壓氣瓶充填高壓氣體，另外四組充填氧化劑並分別供給四顆引擎。因此，每組引擎必須針對不同推力需求提供一定範圍之推力，其中最大推力為60公斤且最大燃燒艙壓為40巴。在最大推力情形下，每組引擎須維持30秒的續航力。以上提及之設計參數會藉由引擎之地面測試取得並進行最佳化。地面測試場所安裝之感測器包含荷重元、壓力計、機械式流量計、差壓式流量計及熱電偶，分別用於推力、壓力、流量及溫度之量測。由於飛行測試階段推力值無法直接藉由感測器量測取得，因此將藉由流量計及壓力計所量測之流量及艙壓值進行推力之估算。推力的控制是藉由控制閥控制進入燃燒艙之氧化劑流量來改變推力大小，此時控制閥亦造成管路壓損變化間接導致艙壓的改變。
首先，藉由地面測試可以獲得引擎燃燒時的氧化劑與燃料用量之比值、氧化劑質量通量與燃料退縮率之關係及燃燒穩定度等性能。地面測試結果顯示，在最大推力60公斤時燃燒艙壓達40巴，如同設計時所設定之參數，且續航力可維持30秒符合需求。在完成全推力及長時間燃燒的研究及分析後，接著測試並分析推力控制之引擎性能，其中包含控制閥角度控制及氧化劑流量控制。地面測試數據顯示，氧化劑流量與推力和燃燒艙壓是成線性關係。這結果驗證了混合式火箭之可控性以及藉由改變氧化劑流量即可進行推力控制之特性。

This research focuses on developing the propulsion system for the Quad Hybrid Rocket Engine Levitating Platform (4-HELP). The 4-HELP is designed to demonstrate and verify the feasibility of developing a levitating and flight platform using hybrid rocket technology. The overall system includes three major subsystems, which are propulsion, structure and avionics. The avionics system include guidance navigation and control, telecommunication and electrical power system. The hybrid rocket engine is applied for the 4-HELP propulsion due to its safety, storable, cost effective and throttling capability characteristics. The equipment of the hybrid rocket propulsion system can be divided into few parts, which are tanks, plumbing and combustion chambers. The nitrogen (N2) is used as the pressurant for pressurizing the 90 wt\% hydrogen peroxide (H2O2) solution used as the oxidizer into the engine to mix with the solid polypropylene (PP) used as the fuel for combustion.

The 4-HELP requires total thrust of 240 kgf with 30 s endurance at maximum mass flow rate of oxidizer, where four engines and six composite tanks of 10 L volume are equipped. Among them, two tanks containing pressurant and four tanks containing oxidizer in order to supply the four engines simultaneously. Therefore, each engine could be operate at the various range of thrust and chamber pressure with the maximum value of 60 kgf and 40 bar respectively. The combustion endurance of each engine is 30 s under the conditions of maximum thrust and maximum oxidizer mass flow rate. To test the performance, this engine is mounted on a thrust stand with load cell, pressure transducers, turbine flow meter, balance flow meter and thermocouple. During a typical flight test, the thrust produced by the engine is not directly obtainable. Therefore, the feedback value of oxidizer mass flow rate and chamber pressure are essential in evaluating the instantaneous performance of the combustion engine during the flight. The thrust is throttled by controlling the oxidizer mass flow rate into the combustion engine using a control valve, where the chamber pressure changes due to the pressure drop created by it.

First, the performances of the designed engine, including the oxidizer to fuel mass ratio, the relationship between oxidizer mass flux and solid fuel regression rate, and the stability of the combustion are investigated at full thrust via hot fire tests. At the state of full oxidizer flow rate, the thrust and chamber pressure of the engine reached 60 kgf and 40 bar respectively as designed and the endurance is able to reach 30 s as required. After investigating the engine performances at full thrust and long combustion duration, the throttling performances of the engine are investigated. The data obtained from the hot fire test results are analyzed, where the chamber pressure and the thrust are found to be proportional to the oxidizer mass flow rate. These results verified that the hybrid rocket engines are capable of throttling simply by controlling the mass flow rate of oxidizer into the combustion chamber.

摘要 i
Abstract iii
Table of Contents v
List of Tables vii
List of Figures ix
Nomenclature xiii
Acknowledgments xviii
Chapter 1 Introduction 1
1.1 Background and Motivation 1
1.1.1 Rocket Propulsion 1
1.1.2 Quad Hybrid Rocket Engine Levitating Platform 2
1.2 Literature Survey 4
1.2.1 Planetary Landers 4
1.2.2 Hybrid Rocket Engines 4
1.2.3 Throttling Abilities of Hybrid Rocket Engines 5
1.3 Objectives 5
Chapter 2 Research Methods 7
2.1 Propulsion System Design and Analysis 7
2.1.1 Engine Design 7
2.1.2 UNIC Simulations 9
2.1.3 Pressure Drop Analysis for Plumbing System 10
2.1.4 Control Algorithm Design for Control Valve 13
2.2 Experimental Facility and Instrumentation 18
2.2.1 General Description 18
2.2.2 Test Instrumentation 18
2.2.3 Test Facility 20
2.3 Test Procedures and Test Conditions 21
Chapter 3 Results and Discussion 23
3.1 Engine Combustion Simulations 23
3.2 Engine Performances 25
3.3 Control Valve Performances 28
3.4 Engine Throttling Capabilities 32
Chapter 4 Conclusion 37
Chapter 5 Recommendations for Future Work 39
References 41
Tables 46
Figures 71

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