臺灣博碩士論文加值系統

在吸氣式高超音速推進引擎中，燃燒室內之燃料噴注方式、燃料與空氣混合效率以及燃料反應過程等問題一直都是很受關注的研究方向。本研究主要對超音速燃燒室流場內進行垂直氫氣噴注這技巧上作出實驗設計，並針對其可自燃性和流場結構進行分析。
對於流場中著火過程的實驗研究，我們利用位於國立成功大學的反射式震波風洞來模擬出超燃沖壓引擎在馬赫六飛行條件下其燃燒室入口的流場環境(條件為M = 2.0, T0=2000K, p0=12MPa, T=1200K, p= 1.5MPa )。對於燃料的噴注(氫氣)，我們設計出一座上游(背向)式爆轟管，並能成功產生達421K的預熱氫氣的噴注。在實驗中，我們在一平板上的一個直徑為2mm的噴孔中將氫氣噴注至超音速流場中。再以高速攝影機和紋影結合的方法對流場進行觀察，亦應用增強式攝影機(ICCD)記錄OH自然螢光影像，以探討其可燃性。
從OH自然螢光影像結果顯示，在氫氣噴流與超音速主流交互作用下產生的邊界層分離之回流區，以及噴流與主流在下游之剪切層內也有明顯的燃燒反應。我們同時在噴注板內安放了兩組熱傳感測器和壓力感測器以提供實驗瞬間之燃燒放熱情況。
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Fundamental to the success of hypersonic air-breathing propulsion system are the efficient injection, mixing, and combustion process that occur inside the combustion chamber. This research is focused on the self-ignitability of a transverse fuel, hydrogen gas, injected into a simulated supersonic combustor environment.
Our experimental approach uses a reflected shock tunnel located in National Cheng Kung University, Taiwan, to provide the high total enthalpy flow associated with the scramjet combustor entry condition of M = 2.0, T0 = 2000 K, p0 = 12 MPa, T = 1200 K and p = 1.5 MPa. A upstream (also called backward) detonation-driven shock tube is also built and used successfully to supply hydrogen fuel at up to T0 = 505 K and p0 = 16.5 MPa for the experiment. In the first set of tests, the ignition of the transverse hydrogen injection from a 2mm wall orifice in a flat plate. In the investigations, a high speed camera is used indicates the location of the shock structures and an intensify CCD (ICCD) detector with hydroxyl (OH) radical wave length filter ( 307 nm) is used to map the regions of ignition in the near-field by the intensity of OH-chemiluminescence. Results show that the significant spontaneous ignition can be clearly observed in the recirculation region caused by the interaction between the traverse hydrogen fuel jet and also in the large scale eddies structure in downstream. Two heat flux sensors and PCB pressure transducer are flush-mounded on the model and also help to give a sign of combustion.

摘要 I
Abstract II
Acknowledgements III
Contents IV
List of Table VI
List of Figures VII
Nomenclature XII
Subscripts XIII

CHAPTER 1 INTRODUCTION 1
1.1 Background and Motivation 1
1.1.1 The Scramjet flight trajectory and combustor entry conditions 3
1.1.2 Types of ground test facilities for supersonic combustion researches 8
1.1.3 Flow-field features of jets in supersonic crossflows 13
1.2 Thesis Objectives 16

CHAPTER 2 REFLECTED SHOCK TUNNEL 17
2.1 Introduction 17
2.2 Theoretical Review and Tunnel Operation 19
2.2.1 Tunnel operation 20
2.2.2 Unsteady gasdynamics in a shock tube 22
2.3 Tunnel Sub-systems 26
2.3.1 Gas supply system 26
2.3.2 Data acquisition system 27
2.4 Tunnel Performance and Calibration 30
2.4.1 Numerical estimations of reservoir and free-stream conditions 30
2.4.2 Measurement of flow properties 36
2.4.3 Shock tube and nozzle performance 38

CHAPTER 3 FUEL INJECTION SYSTEM 48
3.1 Introduction and review 48
3.2 Fuel Injection system design 51
3.2.1 Forward-detonation shock tube design 51
3.2.2 Backward-detonation shock tunnel design 54
3.2.3 Pressure measurement and calibrations 59
3.3 Timing and Synchronization 65

CHAPTER 4 DATA REDUCTION AND FLOW VISUALIZATION TECHNIQUES 67
4.1 Heat Transfer Rate Measurement 67
4.2 Flow Visualization Techniques 71
4.2.1 High-speed Schlieren imaging system 71
4.2.2 OH-Chemiluminescence imaging 72

CHAPTER 5 SUPERSONIC COMBUSTION EXPERIMENTS ON A TRAVERSING JET FROM A FLAT PLATE 74
5.1 Introduction 74
5.2 Test Model 74
5.3 Test Conditions 76
5.4 Experiment Results 78
5.4.1 Heat transfer rate measurements 78
5.4.2 Pressure measurement 81
5.4.3 High speed schlieren images 82
5.4.4 Chemiluminescence images 85

CHAPTER 6 CONCLUSION AND RECOMMENDATION FOR FUTURE WORK 88
6.1 Conclusion 88
6.2 Recommendation for future work 90
6.2.1 The improvement 90
6.2.2 Jet compressibility analysis 90

References 94
Appendix 102

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