In this study, we developed a high-performance and high-thrust hybrid rocket motor using low-melting-point thermoplastic (LT) fuel and swirling oxidizer flow. LT fuel has excellent mechanical and adhesive properties, as well as a high regression rate compared to conventional hybrid rocket fuel. In this study, we conducted several firing tests using swirling oxidizer flow to obtain the fuel regression rate and evaluate its effects on the geometric swirl number (Sg). We determined that the average regression rate of the LT fuel with   = 37.3 was ~2.9 times larger than the axial flow test value. The LT fuel was more susceptible to swirling flow than polypropylene, presumably due to the different physical properties of the fuels. In the swirl flow experiment, we confirmed that the local fuel regression rate behind the fuel is uniform, and it differs from the regression rate seen in the axial flow experiment. For the range of oxygen mass flux values   = 30 − 72,   was fitted to a conventional formula. The results of this fit suggested that the local regression rate at the head region of low-melting-point fuel, such as the LT fuel, cannot be represented only by chemical reactions; therefore, the fluid dynamics of liquefied fuel must be included in the model.

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Our Tokai University Student Rocket Project (TSRP) has launched hybrid rockets since 2003. The attainable maximum altitude of our rockets is about 1 km. The propellants of hybrid rocket are liquid nitrous oxides and wax-based fuel. The hybrid rocket motor is using N 2 O and wax-based fuel. We have 300 N class and 600 N class motors. In August 2012, we launched a rocket which was called "H-28". This paper presents the results of the hybrid rocket motor development and the flight test results. The main goal of this rocket is to launch the rocket toward the sea and recover it successfully after splashdown. Nomenclature e A = nozzle exit area O/F = oxidizer to fuel ratio o A = orifice area a P = atmosphere pressure t A = nozzle throat area c P = combustion chamber pressure d C = discharge coefficient i P = upstream pressure F C = thrust coefficient R = gas constant * th c = theoretical characteristic velocity r = mean regression rate c D = Solid fuel grain outer diameter c T = combustion gas temperature f p D , = port final diameter b t = burning duration i p D , = port initial diameter tank V = oxidizer tank volume F = thrust  = specific heat o G = mean oxidizer mass flux f m  = fuel mass L = solid fuel grain length o m  = oxidizer mass MR = unburned mass fraction * c  = characteristic velocity efficiency f m  = fuel mass flow rate F C  = thrust coefficient efficiency f m = fuel final mass  = circle ratio o m  = oxidizer mass flow rate f  = fuel density p m  = propellant mass flow rate o  = oxidizer density

The method of air-launching a rocket using a launcher suspended from a balloon, referred to as a rockoon, can improve the flight performance of small rockets. However, there have been safety issues and flight trajectory errors due to uncertainty with respect to the launch direction. Air-launch experiments were performed to demonstrate a rail launcher equipped with a control moment gyroscope to actively control the azimuth angle. As a preliminary study, it was suspended via a crane instead of a balloon. The rockets successfully flew along the target azimuth line and impacted the predicted safe area. The elevation angle of the launcher rail exhibited a fluctuation composed of two frequency components. A double-pendulum model with a rigid rod suspended by a wire was proposed to predict this behavior. Significant design parameters and error sources were investigated using this model, revealing the constraining effect of a large mass above the wire and elevation angle fluctuation, which caused trajectory errors due to the friction force on the rail guide and thrust misalignment. Finally, tradeoffs in designing the rail length were found between the launcher clear velocity and elevation fluctuations.

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