在太空科技快速發展的現在，台灣也積極投入太空科技的相關研發工作，而在上層火箭或軌道推進器的開發上，因台灣太空經濟規模較小，且在推力調節與多次點火的需求下，可選擇以混合火箭為目前階段性的推進器發展。
混合火箭機構較為簡單、安全性高，成功大學團隊在HTPB燃料混合火箭相關研究上發展已久，為改善HTPB燃料退縮率較差的情況，研發出與石蠟依比例調配的50P藥柱，並以此發展出30、100、1000以及3000公斤等級推力之發動機。混合火箭之啟動需利用熱源先將固體藥柱熱熔氣化後方可開啟氧化劑控制閥進行噴注點火燃燒，若透過混合火箭自身攜帶之氧化劑進行自發點火，不僅可簡化火箭設計，且能達成重複點火之特性，其中一類點火機制便是將氧化劑先經由觸媒段分解，在將分解過後的高溫產物導入燃燒室內與藥柱自燃點火。過氧化氫原料屬一般工業用途，不受輸出管制、容易取得，本實驗室在長期研究開發下，自主發展出推進級高濃度過氧化氫(High Test Peroxide)精練提純技術，90%過氧化氫絕熱分解溫度更達1067K，因此本研究擬建立一以高濃度過氧化氫為氧化劑經觸媒進行預分解和HTPB/石蠟燃料藥柱自燃點火的混合火箭。
本研究透過推力先決與燃燒室壓力設定的方式進行理論分析進而完成測試用系統之概念設計與參數取得、高負載力觸媒反應室之尺度放大與支撐材填充技術精進改良開發、混合火箭氧化劑通量與藥柱燃面分析建立混合火箭固體燃料幾何構型，成功以此設計準則方法建立可透過自身氧化劑觸媒預分解釋熱點火並持續燃燒產生推力之250N過氧化氫自發點火混合火箭系統，並藉由不同漩流強度之漩渦噴注點火與燃燒特性、發動機性能比較，掌握自發點火混合火箭之設計關鍵。

For the past decades, aerospace exploration has gotten lots of attentions, and the authorities of Taiwan also spent great efforts on autonomous developments of aerospace techniques. For the mission requirements in the future, a low cost, reliable and safety “green” propellant is the first priority.

On the use of upper stage rockets and kick motors, liquid rockets is always a better options because of its high specific impulse, controllable thrust and multi-impulse, but the injection and system design is too complicated and expensive. Without the supports of advanced technologies, it nearly impossible to achieve the goal in the next few years.

Hybrid rockets definitely a good choice to replace liquid rockets with its advantages, like controllable thrust, simpler constructions and high safety. For the last decade, the NCKU hybrid rocket team has succeed launching few N2O hybrid rockets, with thrust level from 100 kgf to 3000 kgf, and our team has also developed a composite silver catalyst used on the H2O2 monopropellant thruster, which has an excellent performance. Therefore, this thesis is going to demonstrate a pre-decomposition H2O2 hybrid rocket using our catalyst’s techniques.

Key words : hybrid rocket, HTP, silver catalyst

INTRODUCTION

For recent years, rocket-grade hydrogen peroxide propulsion systems got a renewed interest due to its low toxicity, low cost and minor impact to the environment. For the hydrogen peroxide concentration over 92% has been called High Test Peroxide(HTP), and the decomposition temperature of 100 wt% HTP can reach 1267K. Therefore, lots of researches was developed on HTP mono-propellant, and silver, manganese based catalyst were commonly used.

Hybrid rocket technology is known for more than 50 years, its separation storage of fuels and oxidizers makes it safer than other propulsion systems. Nowadays, the need for green propellants, safe storability and the use of upper stage rockets made hybrid rockets more attractive. LOX, HTP and N2O are widely used as oxidizers due to their low toxicity and low pollutant characteristics. The initiate of hybrid rocket required a heat source to gasify fuels until reaching a combustible condition, while a ignition device is needed, this makes the construction heavier and more complicated. A pre-decomposition HTP hybrid rocket which can auto-ignite the fuel is developed, and several collages and research units have already gotten tremendous results.

To shorten the period of development, hybrid rockets are definitely a good replacement, our team has lots of experiences on N2O hybrid rockets and also developed a composite silver catalyst using on the HTP mono-propellant, so this thesis is going to demonstrate an auto-ignition hybrid rocket using pre-decomposition HTP.

ANALYST AND DESIGN

The motor design is based on the 2000N thrust requirement of upper stage rockets or kick motors, with 1/8 reduced scale, this thesis is going to verify the preliminary test of a 250 N thrust engine. The CEA (Chemical Equilibrium with Application) code provided a starting point of this investigation. Assumed the combustion chamber pressure of 380 psi and using 90 wt% H2O2 as our oxidizer and 50P (50 wt% of HTPB+50 wt% of paraffin) fuels. From this data we predicted Isp and optimum O/F ratios as showed in figure 1.

The O/F ratio of 7 was selected with the best predicted Isp of 240.6s, compared to the stoichiometric O/F ratio of 7.7, this selected ratio will lead the combustion process to go through a fuel rich combustion, which can prevent the nozzle from corrosion, with the figure we can calculate the other parameters.
m ̇_total=T/(Isp×g)
m ̇_fuel=m ̇_total/(1+O/F)
From the equations above, we can get the mass flow rate of oxidizers and fuels.

Catalyst Bed And Liquid HTP Injector Design
The catalyst bed design is based on the 1N mono-propellant techniques our team has developed, to decompose the H2O2 of 92.75g/s, we used the data captured in the previous experiments, and going through a scale enlargement. The composite catalyst is composed of silver flakes and γ-Al2O3 pellets due to its stronger hardness.

Before entering the catalyst bed, the H2O2 was designed to go through an injector to spread liquid H2O2 uniformly into the catalyst.
m ̇=C_d A√2ρ∆P
The equation above was used for injector design, after assuming the pressure drop of 15 psi and the physical properties, we can estimate the total area required. Therefore, the injector was designed with 16 bores, each bore has a diameter of 1 mm.

Gaseous Injector Design
Injector plays an important role in hybrid rocket mixing mechanism, decomposed H2O2 will be led into combustion chamber, mixing with gaseous fuels which gasified by high temperature gaseous H2O2, the whole process is thermal-related, to reduce the energy loss is our first consideration.
The mass flow rate can be expressed as
m ̇=ρ_2 U_2 A
Where U_2、ρ_2 can be represented as
U_2=√(2γ/((γ-1))×RT_1×[1-(〖p_2/p_1 )〗^((γ-1)/γ)])
ρ_2=ρ_1 〖[p_2/p_1 ]〗^(1/γ)
After decomposing, the exit gas can be regarded as a mixture of O2 and gaseous H2O, assume ideal gas, and set p2/p1 as 0.9, with previous assumption of chamber pressure 380 psi, the pressure ahead of injector can be calculated. We can also calculate the value of density using the ideal gas equation of state, with the assumption of ideal decomposition temperature to be 1067 K and γ=1.26, the ideal exit velocity of 289.106 m/s was estimated. From the calculation, the injector was designed with 8 rectangle grooves each has a length of 5 mm and width of 2 mm.

Fuel Grain
NCKU rocket team has made a great effort on N2O hybrid rocket, and developed the 50P fuel, which composed of 50% paraffin and 50% HTPB, though 50P fuels was chosen as our solid fuels, with the data acquired from previous experiments, a regression rate of 1.5 mm/s was assumed.
m_fuel=ρ×A×L
With the specific weight of 0.9, we can estimate the dimension of the fuel, a hollow cylinder fuel grain was made, with the inner diameter of 20 mm, outside diameter of 53 mm and the length of 180 mm.

Nozzle
From CEA calculations, the total mass flow rate of 106 g/s and the reaction temperature 2700 K during combustion process was known, with γ=1.14
m ̇=p_c A_t γ√(〖[2/(γ+1)]〗^((γ+1)/(γ-1))/√(γRT_c ))
With the equation above, a nozzle throat area of 64.615 mm2 graphite IG-11 was machined as our nozzle.

RESULTS AND DISCUSSION

Characteristics of Composite Silver Catalyst
To verified the feasibility of auto-ignition, the attempt of using pre-decomposition to ignite the fuel was tried. The catalyst was designed and went through a scale enlargement from the 1N monopropellant we developed. Silver flakes and zeolite supports was used, but from the preliminary tests, zeolite was replaced byγ-Al2O3 due to its lack of hardness, and a terrific outcome was obtained. During decomposition process, the temperature over 800K was reached, and the catalyst chamber pressure was maintained steady through whole process, and the structure of the catalyst bed still remain stable after tests.

Combustion Characteristics of 50P Fuels
The motor was successfully auto-ignited in the tests, and from one of the tests listed below, the average thrust of 205 N was measured, the H2O2 flow rate of 92.05g and average chamber pressure of 352.18 psi was also measured in that test, from the figure below, a stable curvature was obtained. And from the definition of specific impulse, a value of 192.12s was reached in case 4.
Isp=F ̅/(m ̇_f+m ̇_o )

Swirl Effects in Combustion Chamber
From the earlier studies, applied swirl injection to the hybrid rocket would increase the performance of combustion efficiency, and the intensity of swirling was defined as the swirl number
S=2/3((1-〖(R_h/R)〗^3)/(1-〖(R_h/R)〗^2 ))tanα
Thus a swirl injector was applied, with different swirl number, the table listed below shows the data measured in the tests, with stronger swirling injection, the combustion efficiency was contrarily decreased, these results differed from the studies, and from the flame observation, the causes was concluded. Due to insufficient gasified fuels, the oxidizer/fuel mixing was poor, the un-gasified fuels was carried to the downstream of combustion chamber.

CONCLUSION
This thesis was tend to demonstrate the auto-ignition of hybrid rocket using decomposition H2O2, a thrust level of 250 N was set, and the swirling injection was applied as well, few results we got was listed below:
Development of high capacity catalyst bed: γ-Al2O3 was used as catalyst support, and zeolite was replaced due to its weaker hardness. During the tests, the catalyst was able to sustain the high H2O2 flow rate and the high temperature.
Establishment of firing system and procedure: This thesis was dedicated in establishing a procedure and method on auto-firing a H2O2 hybrid rocket, and successfully conform the design of H2O2 catalyst bed and hybrid rocket motor.
Succeed auto-firing a hybrid rocket motor: During tests, the time delay of auto-firing are less than 0.08 second, and succeed producing thrust.
The key on promoting the mixing of oxidizer and fuels: In the tests, all combustion conditions were in fuel rich, but from the observation of internal ballistics, a low combustion efficiency condition was observed. With the use of swirling injection, from the previous studies, it should improve the combustion efficiency. In our tests, the regression rate did improved, but the internal ballistics and the flame still shows a low combustion efficiency, this may due to low viscosity of paraffin based fuels, and cause an entrainment effect.

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