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研究生:李孟哲
研究生(外文):Li, Meng-Che
論文名稱:具推力大小與直驅向量控制能力之再生冷卻混合火箭引擎系統開發
論文名稱(外文):Development of a Regeneratively Cooled Hybrid Rocket Engine System with Throttling and Direct-Drive Thrust Vector Control
指導教授:吳宗信吳宗信引用關係
指導教授(外文):Wu, Jong-Shinn
口試委員:陳慶耀廖英皓黃榮芳施聖洋江仲驊
口試委員(外文):Chen, Ching-YaoLiao, Ying-HaoHuang, Rong-FungShy, Shenq-YangChiang, Chung-Hua
口試日期:2022-11-01
學位類別:博士
校院名稱:國立陽明交通大學
系所名稱:機械工程系所
學門:工程學門
學類:機械工程學類
論文種類:學術論文
論文出版年:2022
畢業學年度:111
語文別:英文
論文頁數:137
中文關鍵詞:混合式火箭過氧化氫渦漩注入再生冷卻燃料耗蝕率節流控制推力向量控制
外文關鍵詞:hybrid rocket motorhydrogen peroxideswirling injectionregression rateregenerative coolingthrottlingthrust vector control
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  • 被引用被引用:1
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本論文提出一種具備再生冷卻設計與推力向量控制的混合式火箭引擎系統。再生冷卻設計對於混合式火箭發動機的影響與此混合式火箭引擎的推力可控性皆為本論文的研究重點。首先,本文分別以水平地面靜態測試與數值模擬之方式研究再生冷卻對於採用渦漩注入之混合式火箭引擎的影響與比較。本文使用90%高濃度過氧化氫與聚丙烯分別為液態氧化劑與固態燃料。本文亦提出一種創新的設計將再生冷卻設計概念導入至混合式火箭引擎中。其作用原理為將燃燒艙內所產生的燃燒熱,透過再生冷卻傳導盤及再生冷卻區域內的液態氧化劑進行散熱。除了降低燃燒室外殼溫度以外,也可藉由預熱液態氧化劑而提高氧化劑分解的效率。執行一系列觸媒床、混合式火箭引擎地面點火測試,並完成分析再生冷卻對於混合式火箭引擎在推進表現上的影響。觸媒床地面測試結果顯示,具備再生冷卻設計的觸媒床出口溫度明顯低於未使用再生冷卻設計的觸媒床,但兩者在觸媒床特徵速度上的結果非常接近。表示具備再生冷卻的觸媒床出口溫度雖較未具備者低,其觸媒床效率仍可良好維持。進一步以相同觸媒床設計針對有無再生冷卻設計進行混合式火箭引擎的點火測試,結果顯示具備再生冷卻設計的引擎會有較低的氧燃比及較高的比衝值效率。此外,本論文亦使用一光學雷射3D掃描儀擷取燃燒後的固態燃料表面及量測燃料的縮退率。結果顯示,具備再生冷卻設計的引擎其上下游的燃料縮退率差異較未具備再生冷卻設計小。
針對火箭引擎推力控制,本研究進行一系列水平地面靜態測試藉以了解具備再生冷卻設計的混合式引擎推進性能表現。其中包含以兩種不同且固定的氧化劑流量(425 g/s 及 350 g/s) 進行不同操作時間的點火測試以及在流量區間50~100%的推力大小控制測試。在固定流量為425 g/s與固定噴嘴面積擴張比6.25的情況下的測試結果顯示,此混合式引擎的比衝值高達241.1 s,比衝效率值高達97%且其推力震盪皆小於5%。在推力控制的測試結果中,不管是推力或是艙壓都與氧化劑流量存在著極高的線性關係。其氧燃比也和固定流量的測試結果接近5.6,此氧燃比幾乎不變的優點與過去文獻中曾研究過的混合式引擎相異。總結分析,當推力控制在50~100% (425 g/s)區間時,總比衝與總氧化劑消耗有線性關係,此優點大幅提升混合式引擎對太空發展的應用可能性。
本研究亦提出一計算流體力學模型瞭解混合式引擎燃燒室內的燃燒相關熱流場。首先,藉由引擎地面點火測試的數據作為模擬用的幾何與邊界條件,以比較有無具備再生冷卻設計的影響。此模擬中使用一階化學反應式當作化學模型,透過模擬所得壓力場、速度場、溫度場及渦漩係數了解燃燒室內部情況並與實驗比對。結果顯示,實驗中的燃料耗蝕率與模擬中的渦漩係數有直接的關係。然而,由於有無具備再生冷卻設計的測試結果中,兩者的氧燃比與殘餘燃料內徑尺寸不同,而此皆會影響數值計算的結果。因此,本研究另在固定氧燃比及固定殘餘燃料內徑尺寸的條件下,以減少變因、簡化數值分析並透過調整不同上下游燃料縮退率差異來了解燃料上下游耗蝕率差異對燃燒室的影響。結果顯示,燃燒室內燃料剖面可藉由渦漩係數切分成兩種縮退率區域: 非典型區與典型區。在非典型區中,燃料縮退率主要受因渦漩擴張效應產生的徑向衝擊影響。此外,在此區域中不同的上下游燃料縮退率差異都會有極接近渦漩係數計算結果。在典型區中,由於上下游燃料縮退率差異較大者在燃燒室上游處減少的切線動量較多,導致其渦流係數較差異較小者小,代表若上下游燃料縮退率差異越小,則在燃燒室中下游的渦漩係數越高,其渦漩強度較高,就會有較高的燃料縮退率,此也使具再生冷卻的引擎具有較低的氧燃比、較高的混合效率及較高的比衝效率值。本論文更進一步延伸使用數值模型研究燃料內徑尺寸與推力控制對渦漩注入的影響。計算結果顯示渦流強度會隨著內徑增加而增加,尤其是在燃料中下游的典型區,此現象亦與點火測試結果相同:隨著燃燒時間增加,氧燃比會稍為下降且比衝效率值會上升。另一方面,推力控制時渦漩強度並未隨著氧化劑流量調整而改變,此也解釋了為何在推力控制點火測試中氧燃比並未隨著流量變動而明顯改變。
在充分了解本引擎的性能後,本研究藉由一商用直驅式電動馬達驅動此引擎當作動力源進行推力向量控制測試,並在一座垂直測試台上進行實驗。此實驗中關於推進表現及推力向量控制表現的數據皆由壓力計、流量計及角度計等量測得到。結果顯示,推力的向量控制相當精準,誤差低於0.1度。HTTP-3A第二節箭體中的推進次系統是使用共四組引擎及直驅式電動馬達組合而成,而此推進次系統已接連完成短時間、長時間垂直點火測試、懸浮測試及第二節飛行測試。在幾次系統整合測試中,引擎推力控制表現優異,其精準度及震盪性皆低於5%。在推力向量方面,向量角度的精準度及震盪性皆小於0.1 度及0.05 度。尤其在飛行測試中,即便不斷有側風干擾的情況下,兩者優異的表現讓箭體可以平穩升空,此也無疑是混合式火箭在太空技術發展上的重大突破。
This thesis proposes the research and the development of a regeneratively cooled (RC hereafter) hybrid rocket engine system with capacity of throttling and thrust vector control that using swirled 90%-wt hydrogen peroxide and polypropylene as oxidizer and fuel, respectively. Firstly, the experiments for swirling-type hybrid rocket motors without and with the design of regenerative cooling were performed and compared. A series of ground hot-fire tests including catalytic-bed only and combustion chamber were performed to obtain the effect of regenerative cooling on the performance of hybrid rocket motor. The new idea by implementation of regenerative cooling using the oxidizer resulted in a corresponding temperature drop at the catalytic bed outlet. The resulting C* efficiency of the catalytic-bed is comparable with the one without RC. For hot-fire tests, the results showed that the case with RC had a better engine Isp efficiency with a lower O/F ratio as compared with the one without RC. In addition, the topologies of the burned fuel grains measured by an optical 3D scanner showed that the variation of regression rate between the upstream and downstream fuel was smaller for the case with RC than the other.
Ground static hot-fire tests with two fixed m ̇_ox of 425 g/s (100%) and 350 g/s (82%), and throttled m ̇_ox with the same nozzle with an area expansion ratio of 6.25 were performed to examine the engine’s propulsion corresponding performance. The hybrid rocket motor achieved an impressive sea-level I_sp to 241.1 s, equivalent to an I_sp efficiency of more than 97%, and a thrust uncertainty of less than 5% for fixed m ̇_ox of 425 g/s. For throttling tests, the thrust and chamber pressure were both found nearly proportional to the m ̇_ox with different proportionalities. The overall averaged O/F ratio of the tests was 5.61 which is hardly shifted unlike most test data of hybrid rocket reported previously. In brief summary, the total impulse was linear to total oxidizer mass consumption in the range of 50-100% of 425 g/s.
A CFD model was built to investigate the combustion related phenomenon inside the proposed hybrid combustion chamber. First, the experimentally measured regression rates at different fuel grain portions were set as the boundary conditions of the fuel inlet. A single-step global chemical reaction was assumed in the model. The solutions of pressure, velocity, temperature, and swirling number were investigated and compared with the experimental results wherever possible. The calculated distribution of the swirling number was found to correlate reasonably well with the experimentally observed regression rate from the hot-fire test. Two distinct regression-rate regions, named Non-classical and Classical, which could also be observed in the hot-fire tests. In addition, a series of simplified simulations with a fixed O/F ratio, a fixed port size, and different regression rates between upstream and downstream fuel were performed to investigate the effect of the variation of the regression rates, fuel port size, and throttling on the swirling combustion chamber. For the effect of the uniformity of regression rate, the results showed that the fuel regression was dominated by the gas impingement due to the swirling gas expansion effect in non-classical region. For the rest of chamber, the case with a more uniform regression had a larger swirling intensity than others. This explained why the regeneratively cooled hybrid motor performed with a lower O/F, a better mixing efficiency, and a higher engine I_sp efficiency than the other. Furthermore, the results of different port sizes showed that the larger port size had larger swirling intensities particularly for the classical region, which could also be observed from the hot-fire test results that longer operation had a higher I_sp efficiency and lower O/F. In addition, the calculated swirling intensity rarely changed during throttling, which also explains why there was nearly no O/F shift in the hot-fire test.
Finally, the hybrid motor was driven by a commercial directly-drive electrical motor (DDMotor hereafter) for controlling its thrust direction. A vertical hot-fire test stand was built to conduct the thrust vector control hot-fire tests. The thrust vector was successfully controlled with an accuracy of less than 0.1o. There were total four sets of hybrid motors and DDmotors integrated in the propulsion system with TVC for the 2nd stage of HTTP-3A Hybrid Rocket. The propulsion subsystem was validated using vertical hot-fire tests, hovering flight tests, and realistic flight tests. The results showed that the throttling performance was impressive with an accuracy of less than 0.1o and an oscillation both less than 5% for all tests. The performance of the TVC from the above-mentioned tests was demonstrated with an accuracy and an oscillation less than 0.1o and 0.05o, respectively. Finally, the proposed hybrid propulsion system demonstrated an important milestone for hybrid rocket development in future application in space technology.
Table of Contents
Acknowledgements i
摘要 ii
Abstract v
Table of Contents viii
List of Figures xi
List of Tables xvi
Nomenclature xvii
Chapter 1 Introduction 1
1.1 Background and Motivation 1
1.2 Literature Survey of Development of Hybrid Rocket Motor 2
1.2.1 Hybrid rocket motor with swirl injection of the oxidizer 2
1.2.2 Solid fuel regression rate and O/F ratio shift 3
1.2.3 Applications of high density hydrogen peroxide on space technology 5
1.2.4 Capacity of throttling the hybrid rocket motor 6
1.2.5 Applications of regenerative cooling on liquid rocket engine 8
1.2.6 Numerical investigation of modeling a hybrid rocket motor 10
1.2.7 Conventional thrust vector control system 11
1.3 Specific Objectives and Roadmap of This Dissertation 12
Chapter 2 Research Methods 13
2.1 120-kgf Hybrid Rocket Motor Design 13
2.1.1 Throttling capacity of the hybrid rocket motor 13
2.1.2 120-kgf hybrid rocket motor design without the regenerating cooling 15
2.1.3 120-kgf regeneratively cooled hybrid rocket motor design 17
2.2 Experimental Test Facility and Instrumentation 18
2.2.1 Propellant feed system 18
2.2.2 Data acquisition system and sensors 19
2.2.3 Main throttle valve and balance flow meter 20
2.2.4 Thrust vector control system design 20
2.3 Hot-Fire Test Matrix 21
2.3.1 Effect of regenerative cooling design on HRM 21
2.3.2 Propulsive investigation of the regeneratively cooled HRM 22
2.3.3 Vertical thrust vector control (TVC) hot-fire test 23
2.3.4 HTTP-3A 2nd stage system 30-s and 60-s vertical test 23
2.3.5 S2 hovering flight test & high-altitude flight test 24
2.4 Numerical Investigation Method 25
2.4.1 Turbulent model for swirled hybrid rocket motor 25
2.4.2 Simplified reaction model for hybrid combustion chamber 28
2.4.3 Computational fluid domain and geometry of combustion chamber 30
2.4.4 Boundary conditions settings based on hot-fire test results 30
2.4.5 Simplified numerical investigation of deviation of regression rate 31
2.4.6 Effect of fuel port size and throttling on calculated swirling number 31
Chapter 3 Propulsive Performance of RC HRM 33
3.1 General Description 33
3.2 Effect of Regenerative Cooling Design on HRM 33
3.3 3D Topology and Regression Rate Analysis of Solid Fuel Grain 34
3.4 Single Step of 425 g/s Baseline Hot-Fire Test Results 35
3.5 Single Step of 350 g/s Baseline Hot-Fire Test Results 37
3.6 Multiple Step Hot-Fire Test Results 38
Chapter 4 Numerical Results of Swirled HRM Model 40
4.1 General Description 40
4.2 Grid and Time-Step Convergence tests 40
4.3 Effect of Regenerative Cooling in Numerical Investigation 41
4.4 Effect of Deviation of Regression Rate on Swirling Intensity 42
4.5 Effect of Fuel Port Size on Swirling Intensity 43
4.6 Effect of Throttling on Swirling Intensity 45
Chapter 5 Integration Test Results of HTTP-3A 2nd Stage 46
5.1 General Description 46
5.2 Single Vertical Thrust Vector Control Test 46
5.3 S2 30-s and 60-s Vertical Hot-Fire Test 47
5.4 S2 25-s Hovering Test & High-Altitude Flight Test 48
Chapter 6 Conclusions and Recommendations for Future Work 50
6.1 The effect of the regenerative cooling on a hybrid rocket motor 50
6.2 The performance of the RC HRM and its throttling capacity 52
6.3 120-kgf HRE system with throttling and TVC 54
6.4 Recommendations for Future Work 55
Reference 56
List of Publications 64
Journal Papers 64
Conference Papers 65
Figures 67
Tables 121
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