堀 恵一

The third Epsilon launch vehicle, which was the Enhanced Epsilon launch vehicle optional configuration, was successfully launched with the payload of ASNARO-2 in January 2018. The post flight analysis was conducted. The chamber pressure and thrust include residual thrust of three main motors and chamber pressure of SMSJ and SPM were analysed to confirm the effectiveness of design and production methodologies. All solid propulsion systems for the third Epsilon launch vehicle showed a very good behaviour during the flight.

Since 2000, Japan Aerospace Exploration Agency (JAXA) is involved in the research and development (R&D) of hydroxYlammonium nitrate (HAN)-based liquid monopropellant, SHP163 which is composed of HAN, ammonium nitrate (AN), methanol and water for its possible utilization as an alternative of hydrazine toxic component. The freezing point of SHP163 is much lower, the density and the density-Isp are higher than those of hydrazine. The toxicity of SHP163 is also lower than hydrazine, therefore easy handling. Moreover, SHP163 monopropellant showed stable burning rate measured by strand burner and its thermal decomposition of HAN-based occurred in two steps. However, the catalytic process showed a violent exothermic reaction of HAN-based in one step at low temperature onset around 65 C. The application of SHP163 on thruster firing tests showed that pressure slopes have been increased in the case of honeycomb catalysts and burning reactions have been improved by new synthesized Ir-CuO honeycomb catalyst. The obtained results on thruster 20 N type demonstrate that SHP163 is a good candidate as green propellant and proved the reason why it has been selected to be used as hydrazine alternative monopropellant for "Innovative Satellite Technology" project launched in 2016 by JAXA, Japan. (C) 2016 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

The development of Enhanced Epsilon is mainly the renewal of the second stage, and also includes each subsystem's improvement. The second-stage motor M-35 was newly designed and manufactured. In order to verify the design, the static firing test of the M-35 under the condition of vacuum ambient was conducted in 2015. The Epsilon's second flight, which was the first Enhanced Epsilon launch vehicle, was successfully conducted with the payload of Exploration of energization and Radiation in Geo-space (ERG) in Dec., 2016. After the flight, chamber pressure and thrust include residual thrust were analysed to confirm the effectiveness of design and production methodologies. All solid propulsion systems for the Enhanced Epsilon launch vehicle showed a very good behaviour during the flight.

Basic properties such as toxicity, thermal decomposition temperature, burning rate, and detonability for hydroxyl ammonium nitrate (HAN)/ ammonium nitrate (AN)/ methanol/water mixture solutions are discussed in the paper based on experimental data and qualitative evaluation. From toxicity evaluation, it is expected that HAN based monopropellants have potential as a "Green" propellants, which means we can treat them without self-contained atmospheric protective ensemble (SCAPE) suits. Thermal decomposition temperatures of HAN/AN/methanol/water mixture solutions with various mixture ratios have also been evaluated. Burning rate measurements of combustion waves under pressurized conditions have been taken for many IIAN/AN/methanol/water mixture propellants in order to understand the burning rate and process. Then, the contribution of each component of the mixture has become clear. Finally, we found that stoichiometric solution (SHP163) has a low burning rate over a wide range of pressures. From the large scale gap test (LSGT). we found that SHP163 has low sensitivity compared to other high energetic materials.

The Epsilon launch vehicle, the newest version of Japan's solid propulsion rocket, made its maiden ight in September of 2013. The purpose of the Epsilon rocket is to provide small satellites with responsive launching with low-cost, user-friendly and efficient launch system. Now that the rst ight was successfully nished, JAXA has been conducting intensive researches on a next generation Epsilon to launch a more powerful and lower cost version of Epsilon (Evolved Epsilon). In order to minimize technical risks and to keep up with demand of future payloads, JAXA plans to take a step-by-step approach toward Evolved Epsilon. As the first upgrade toward Evolved Epsilon, JAXA has started the development of Enhanced Epsilon. The Enhanced Epsilon is required to enhance launch capability by ERG satellite mission which have decided to change an orbit to be put into to farther and also to improve on-board capability in size and in weight by ASNARO2 satellite mission. The development of Enhanced Epsilon is mainly the renewal of the second stage, and also includes the each subsystem's improvement. The main change of the solid propulsion system is exposure of the second motor M-35. Currently, the design has been finished. The outside diameter of the motor case is expanded into approximately 2.5 m in order to increase the amount of the solid propellant and the outer shell of the motor case is used as the outer shell of the launch vehicle. Solid propellant which can the high-performance equal to a conventional upper-stage motor developed newly, reducing the cost. A general front ignition system is adopted instead of the rear ignition system of the throw-away type which was adopted for the previous motor. A new development material is applied to the case lining. An expansible nozzle is not adopted because compatibility of high-performance and cost reduction. In order to verify the design, the static firing test of the second motor M-35 on condition of vacuum has been conducted. This paper describes overview of development of the solid propulsion system for Enhanced Epsilon and the results of the M-35 static firing test.

HAN-based liquid monopropellant has been studied for its possible substitution of hydrazine toxic component. The onset temperature of decomposition was provided by DTA-TG thermal analysis. Moreover, the effect of 5M HNO3 and 5M NH2OH addition on catalytic decomposition was also examined. The major products measured by mass spectrometer were N2, NO, N2O, NO2, H2O and NH3. Otherwise, the burning rates of HAN-based monopropellants were measured from the strand burner videos taken by high speed camera. The effect of methanol addition; as fuel; on the burning rates was demonstrated. On the other hand, HAN-based liquid monopropellant was tested and decomposed in 20N thruster and different catalysts were compared. Honeycomb catalysts were tested instead Shell 405 catalysts, the obtained data demonstrate the improvement of burning reactions and pressure slopes after HAN-based decomposition.

This paper reports on the thermal and combustion behaviors of ammonium dinitramide (ADN). The thermal behavior is measured by a pressure thermogravimetric analysis (TGA) at pressures below 8MPa. The burning rates of pure ADN and ADN/ammonium nitrate (AN) mixtures are measured in the range 0.2-12MPa, and the burning temperature profiles are obtained using thermocouples with diameters of 5 and 25m. This report mainly focuses on the condensed-phase behavior in the vicinity of a burning surface. The temperature profiles are complicated because the ADN decomposition and AN dissociation compete during the condensed phase, and the bubbles of the decomposition gas and gas-phase flame also affect the surface temperature. AN addition helps to understand the effects of AN during the condensed phase, and it was shown that the burning temperature rises to the critical temperature of AN. Based on these experimental results, the pressure dependency of the burning rates is also discussed.

The influence of the catalyst support shapes on the performance of liquid monopropellant thrusters was investigated. In the present work, two supports: monolith honeycombs and alumina grains were tested and their relative performances were compared. Ir-monolith catalysts with total metallic contents of 30 wt% in mass were prepared by the wet impregnation method. The characterization of these materials, before and after their use for liquid monopropellant decomposition in a satellite thruster, was performed by measurement techniques of specific surface area, H-2 chemisorption, ICP, SEM and TEM measurements. The catalytic decomposition tests display the best catalytic activity for the grain catalyst (Shell 405) with a complete decomposition of liquid monopropellant, due to a good contact between solid-liquid phases, and monolith catalyst with longer catalyst bed. Moreover, the monolith catalyst leads only to partial liquid monopropellant decomposition, due to injection model and preferential channels through the catalyst bed. The performance of the thruster when using monolith honeycomb and alumina grains as the catalyst bed was evaluated by measuring the product-gas temperatures and pressures at different points of the catalyst bed.

In order to elucidate the mechanism of extremely high burning rates of some Hydroxylammonium Nitrate (HAN) based solutions, the combustion characteristics of HAN aqueous solutions were studied. It was found that the role of the two-phase region is very important and the intense boiling of water by superheat is responsible for the high burning rate. The evaluated bubble nucleation rate coincides with burning rate at very high region. Further, the combustion mechanism of propellant solution is discussed. Hydrodynamic instability was taken into account and the pressure dependency of the instability was estimated. The estimation of hydrodynamic instability supports the observed phenomena. It was found that the instability is strongly affected by Markstein number.

Hybrid rockets using Glycidyl Azide Polymer (GAP) as a solid fuel has been studied. Linear burning rate spectrum of GAP was spread with a dilution of polyethylene glycol (PEG), and basically, self combustible mixtures are used for the gas hybrid rocket motor, and non-self combustible mixtures for traditional hybrid motor. Ultrasonic measurement is employed using three sensors especially developed for this use to evaluate the instantaneous surface regression rates in the traditional hybrid motor. Results from three sensors set at the upstream, middle and downstream of 1.3m length motor give indispensable information for the establishment of the flame model of hybrid motor.

Hybrid rockets using Glycidyl Azide Polymer (GAP) as a solid fuel has been studied. Linear burning rate spectrum of GAP was spread with a dilution of polyethylene glycol (PEG), and basically, self combustible mixtures are used for the gas hybrid rocket motor, and non-self combustible mixtures for traditional hybrid motor. Ultrasonic measurement is employed using three sensors especially developed for this use to evaluate the instantaneous surface regression rates in the traditional hybrid motor. Results from three sensors set at the upstream, middle and downstream of 1.3m length motor give indispensable information for the establishment of the flame model of hybrid motor.

A basic study to clarify the combustion mechanism of glycidyl azide polymer (GAP) has been conducted. Temperature during the strand burner and 60-mm diameter motor tests was measured. The strand tests were performed with 2.5-m diameter S-type thermocouple, embedded in GAP samples, with pressure ranging from 1 to 10 MPa. The 60-mm diameter motor tests were done with end-burning grains and the temperature inside the motor was measured with a 1.0 mm diameter K-type thermocouple with a pressure range from 3 to 10 MPa. The motor tests show the gas temperatures to be approximately 80 K higher than the strand tests and both temperatures are significantly lower than adiabatic temperature. The efficiency of C*, C*, is in the range of 0.7 to 0.85 depending on pressure and L*. Combustion residue of GAP was investigated and it was found to be composed of soot (black in color), high viscosity residue, and a yellow powder, which was only observed at high pressures. These residues were analyzed by means of Scanning electron microscope (SEM) and Fourier transform infrared spectrometer (FTIR), and mass balance was also measured. One-dimensional three-phase mode combustion model of GAP has been constructed based on the Beckstead model. Modifications were made taking into account experimental observations. A blow-off mechanism was added in residue behavior and full kinetics chemistry was entrained in the bubbles at the two phase region. The burning rate and temperature profile were numerically simulated adjusting for kinetic parameters. The rapid temperature increase and final temperature are expressed well in this simulation and the calculated burning rate coincides well at medium pressures.

Currently, the market of airbag inflators as automobile safety devices is still expanding and the key technology in airbag inflation depends on the performance of a gas generant inside the inflator. The gas generant has been developed for modifying inflator performance for airbag deployment in automobiles, usually such a target would be higher gas mol per weight of gas generant, lower gas temperature, and more filterable slug formation of residue inside the inflator. One representative composition of only fuel and oxidizer as the gas generant, which satisfies the target performance, was selected and studied in order to prevent the reaction mechanism from burning the inflator in the airbag system. For a basic study of recent typical gas generants, an initial reaction model with input data was constructed from many experimental data, including the burn rate, burn temperature, SEM image of residue, high-speed video camera image of gas generant burning, and high-pressure differential thermal analysis (DTA).

In this study, reactive B/Ti multilayer igniters were investigated for the noncontact ignition of a micro solid rocket array thruster in vacuum. When current is supplied to the B/Ti multilayer igniter, the chemical reaction: 2B + Ti -> TiB2 + 1320 cal/g occurs, and sparkles are spread to a distance of several millimeters or more. The B/Ti multilayer igniters with three sizes were fabricated, and tested in six configurations of solid propellant. Although one rocket with ignition charge was ignited successfully, the noncontact ignition of the solid propellant was not achieved. However, the B/Ti multilayer igniters themselves generated small impulses of 10(-6) Ns order, suggesting the possibility of self-propulsion. (C) 2008 Elsevier B.V All rights reserved.

In this study, reactive B/Ti multilayer igniters were investigated for the noncontact ignition of a micro solid rocket array thruster in vacuum. When current is supplied to the B/Ti multilayer igniter, the chemical reaction: 2B + Ti -> TiB2 + 1320 cal/g occurs, and high temperature plasma is discharged to a distance of several millimeters or more. The B/Ti multilayer igniters with 3 sizes were fabricated, and tested in 6 configurations of solid propellant. Although one rocket with ignition charge was ignited successfully, the noncontact ignition of the solid propellant was not achieved.

The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3–4 W with RK and 4–6 W without RK. Measured impulse thrusts were from 2×10⁻⁵ Ns to 3×10⁻⁴ Ns after the calculation of compensation for air dumping.