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.
Hybrid rocket motor/engine forms a boundary layer combustion in the chamber because it is composed of liquid oxidizer and solid fuel. Liquefying polymers like paraffin are recently employed as hybrid rocket fuel due to higher regression rate compared to conventional fuels. The observation studies of combustion flame on liquefying fuel surface were conducted at low oxidizer mass flux. And its combustion detailed process is still unclear. Therefore, to examine the melting fuel behavior, a two-dimensional chamber with observation window was performed at GOX 104 kg/m2s and PC 2.0 MPa. In addition, the temperature profiles near the fuel surface were measured using a thermocouple. Distance between fuel surface and flame zone affected by GOX was obtained using thermocouple technique.
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 Sg = 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 Goxlo = 30–72, ṙloave 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.
The development of a direct injection gas-hybrid rocket system using glycidyl azide polymer (GAP) as a solid fuel for the thrusters of very small satellites is described. Furthermore, a performance evaluation and the combustion characteristics of the propulsion system are presented. GAP is capable of self-decomposition and generates fuel-rich gas, which makes it viable as a fuel gas-hybrid power source for rockets. GAP also has a higher density compared to other polymers such as hydroxyl-terminated polybutadiene (HTPB), and the high-density specific impulse enables the development of a small thruster system. Gaseous oxygen was used as the oxidizer for the first test of the gas-hybrid rocket. The gas generator was tested using a 60 mm diameter motor with an end-burning GAP grain. The experimental combustion pressure was initially set at 1 MPa, and adjustments to the oxygen flow were made based on the test results of the gas-generator combustion. The resulting ignition smoothness and combustion stability were observed. Excellent characteristic velocity efficiency (90%)—larger than that of a classical hybrid rocket motor-was obtained. Moreover, quenching of the GAP gas generator was achieved after the oxidizer injection was stopped, which implies that this system has the capability of re-ignition.
Preliminary experimental studies on the flash pyrolysis behavior of low-melting-temperature thermoplastic (LT) were conducted under typical hybrid rocket operation conditions to obtain the decomposition characteristics of the fuel. LT fuel is a paraffin-added thermoplastic elastomer used in hybrid rocket fuel or solid propellant binders. The temperature profile at or near the surface was measured at 2 MPa chamber pressure and 50 kg m-2 s-1 oxidizer mass flux by a 25 μm thermocouple to estimate the phase structure of the fuel. The paraffin oil was flash pyrolyzed in a pyrolysis temperature range of 758 K to 1,313 K (maximum heating rate: 6,400 K s-1) with a gas chromatography mass spectrometer. Under each temperature condition, the paraffin oil produced a unique pyrolysis mass-spectrometry spectrum. In high-temperature regions, the mass spectra indicate lower molecular weight-range products. Benzene, methylbenzene, and vinylbenzene were obtained as pyrolysis products from the paraffin oil at a pyrolysis temperature of 1,037 K. These results suggest that the formation of aromatic compounds dominated the paraffin-oil pyrolysis process. The pyrolysis behavior of LT fuel was observed by combining the results of the LT-fuel temperature profile and the pyrolysis process in paraffin oil. The result shows that decomposing the LT fuel may form aromatic compounds around the burning surface.
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 N2O 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.
