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Mitsuaki TANABE, Masanori SAITO, W. Zach HALLUM, Eric J. MEIER, Trista ...
2016 Volume 14 Issue ists30 Pages
Pa_1-Pa_6
Published: 2016
Released on J-STAGE: February 16, 2016
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Self-excited combustion oscillations in a model rocket combustor is investigated experimentally. A unique dump combustor, CVRC (Continuously Variable Resonance Combustor), is employed to realize a well-controlled self-excitation. The combustor has a coaxial injector whose oxidizer post has a choked inlet that is variable in length allowing for a desired response for the self-excitation. Gaseous methane and decomposed hydrogen peroxide are supplied and burnt in an optically transparent combustor. The flame inside the combustor during hard oscillation is observed by high-speed (20 kfps) CH*-band emission imaging. Together with the images, pressure fluctuations near the dump wall are recorded. As a result, the existence of a nonlinear acoustic wave (N-wave) is suggested when the amplitude of the pressure oscillation exceeds roughly one tenth of the mean pressure. The relation between the occurrence of N-wave and the CH*-band emission oscillation is investigated by applying snapshot POD (Proper Orthogonal Decomposition). Particular spatial modes of the flame emission oscillation are found to appear in accordance with the occurrence of the N-wave.
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Nobuaki SUGIU, Kazunori MOTOHASHI, Masanori SAITO, Mitsuaki TANABE
2016 Volume 14 Issue ists30 Pages
Pa_7-Pa_12
Published: 2016
Released on J-STAGE: May 26, 2016
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The flame curvature and the lift-off height of a triple flame under acoustic oscillations are investigated. The multi-slot burner, which can make uniform streamwise flow velocity and stabilize the triple flame two-dimensionally is employed. The triple flame is formed at the anti-node of velocity oscillations. The flame curvature and the lift-off height of the triple flame are measured from images that are taken by a high speed camera. The fuel concentration gradient of the mixing layer which flows into the flame surface with acoustic oscillations is calculated by using the model of the meandering mixing layer. As a result, the flame curvature changes periodically and its period of change corresponds with that of acoustic oscillations approximately. When the triple flame moves in the direction of the air flow, the flame curvature increases. When the triple flame moves in the direction of the mixture flow, the flame curvature decreases. It is estimated that the fuel concentration gradient changes in the half-period of acoustic oscillations. However, the flame curvature changes in the same period of acoustic oscillations.
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Simona SILVESTRI, Maria P. CELANO, Christoph KIRCHBERGER, Gregor SCHLI ...
2016 Volume 14 Issue ists30 Pages
Pa_13-Pa_20
Published: 2016
Released on J-STAGE: December 09, 2016
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In the current study the effects of oxidizer post recess length variation in a shear coaxial injector have been experimentally investigated. Different injector configurations are used to inject oxygen and methane, both in gaseous form, into the combustion chamber at pressure levels between 10 and 20 bar. It has been observed that the GOX post recess enhances the mixing between the propellants when its length is longer than one GOX post exit diameter. The pressure drop across the injector increases with the recessed oxygen tube compared with the flush mounted case. The variation in wall temperature and pressure axial profile, the pressure drop at the injector and the influence of the injector setup on the heat loads to the wall are the focus of the present investigation.
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Yuki MORI, Shunya SHIMADA, Shuhei TAKAHASHI, Tadayoshi IHARA, Shinji N ...
2016 Volume 14 Issue ists30 Pages
Pa_21-Pa_26
Published: 2016
Released on J-STAGE: September 30, 2016
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Two-band emission method for measuring temperature distribution using CO2 or H2O as the emission medium was developed. In the calibration test using McKenna hydrogen flat flame burner, the developed method had similar quantitativity compared with the other measurement methods, CARS and SiC radiant method. When we applied the developed two-band emission method to high enthalpy supersonic wind tunnel experiment, we could capture the behavior of shockwaves and expansion fans in the supersonic flow and the reconstructed temperature distribution was reasonably quantitative.
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Yu-ichiro IZATO, Hiroto HABU, Atsumi MIYAKE
2016 Volume 14 Issue ists30 Pages
Pa_27-Pa_30
Published: 2016
Released on J-STAGE: November 26, 2016
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The condensed phase decomposition reactions of ADN were investigated both experimentally and theoretically. Thermogravimetric-differential thermal analysis coupled with mass spectrometry (TG-DTA-MS) was employed to generate Friedman plots for the thermal decomposition of ADN with the evolution of N2O and N2. The activation energy associated with the evolution of N2O during initial decomposition was found to be 150 kJ/mol. Chemical equilibrium calculations based on the reaction N(NO2)2- + NH4+ ⇌ HN(NO2)2 + NH3 demonstrated that the concentration of HN(NO2)2 gradually increased with temperature, although the HN(NO2)2 to N(NO2)2- ratio was still only approximately 3.1 × 10-6, even at the decomposition temperature of 130°C. Thus, molten ADN was found to contain primarily N(NO2)2 and NH4+ with only minor amounts of liquid HN(NO2)2 and NH3. The reaction ADN → N2O + NH4NO3 was also investigated using ab-initio calculations at the CBS-QB3//ωB97XD/6-311++G(d,p) level. It was determined that four reaction pathways are possible via different transition states. The energy barrier of 161 kJ/mol obtained from these calculations agreed with the experimental value.
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Kazuya IWATA, Shinji NAKAYA, Mitsuhiro TSUE
2016 Volume 14 Issue ists30 Pages
Pa_31-Pa_38
Published: 2016
Released on J-STAGE: November 26, 2016
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Oblique detonation under incompletely premixed conditions has not been well understood and is of great concern when difficulty of high-speed premixing in Oblique Detonation Wave Engine (ODWE), which is one of the most potential hypersonic aerospace propulsion systems, is taken into account. This study numerically investigated effects of fuel concentration gradients on oblique detonation and shock-induced combustion formed on a 28.20° wedge by solving two-dimensional Navier-Stokes equations with a detailed chemical kinetic mechanism of hydrogen-air combustion. Oblique detonation with smooth-transition formed at a Mach number of 8.00, a static temperature of 300 K, and a static pressure of 8.50 kPa was referred as the completely premixed case. Fuel concentration gradients were described by the Gaussian function. At the maximum equivalence ratio of 2.00, Smooth-transition was replaced by abrupt-transition. When maximum equivalence ratio exceeded 3.00, a V-shaped flame front appeared with its leading edge located away from the wedge, which caused two separate triple-points to be observed. Second triple point appeared at the intersection of the incident shock or the detonation front and a reflected shock generated by compression waves on the lower side of the deflagration front. Increase of the front angle enabled intensive combustion to be maintained downstream of it.
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Shota KAMEYAMA, Shun NAWATA, Tsunetaro HIMONO, Keita KIKUCHI, Masashi ...
2016 Volume 14 Issue ists30 Pages
Pa_39-Pa_44
Published: 2016
Released on J-STAGE: November 26, 2016
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To achieve stable detonation wave propagation to large-bore pulse detonation engine (PDE) combustors, we investigated an initiator for PDEs that uses a pre-detonator, reflector, and driver gas. In this initiator, a planar detonation wave from the pre-detonator becomes a cylindrical detonation wave after collision with the reflector. Wakita et al. previously posited two hypotheses regarding the dominant factors that determine the threshold of propagation to the target gas, which is a stoichiometric hydrogen–oxygen mixture diluted with nitrogen. This study reveals whether the threshold is determined by w/λ or λ/r. To analyze the effect of channel width w on the transition of the cylindrical detonation wave, experiments were conducted for w = 10 mm and w = 15 mm. However, the results could not elucidate whether the threshold is determined by λ/r or w/λ. To clearly distinguish between the effects of w/λ and λ/r, a narrow channel width w' = 3 mm was chosen and implemented using a torus-shaped obstacle. The results showed that a cylindrical detonation wave propagates without quenching when the nitrogen concentration is above 40%, corresponding to the cell size λ greater than w. Accordingly, the cylindrical detonation propagation threshold was determined to be independent of w/λ.
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Daiki MUTO, Hiroshi TERASHIMA, Nobuyuki TSUBOI
2016 Volume 14 Issue ists30 Pages
Pa_45-Pa_52
Published: 2016
Released on J-STAGE: November 26, 2016
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Effects of injector geometries on cryogenic co-flowing planar jets under a supercritical pressure are numerically investigated. The present study covers a wide range of injector exit geometries which focuses post lip height and recess length, and evaluates these effects on mixing characteristics. A hybrid ILES/RANS methodology is applied to simulate wall-bounded injector regions. The results show that thicker post lips generate larger vortices behind the post lip, resulting the comb-like structure of the rolled-up inner dense jet. As a result, the mixing is well improved, and the inner jet potential core is shortened. The recessed injectors additionally induce a large-scale flapping motion of the inner jet and further enhance the mixing. The frequency analysis with velocity fluctuations demonstrates that the vortex shedding behind the post lip has a frequency which depends on the post lip height. The recessed injectors induce another low-frequency peak of the flapping motion, and which value is independent of the post lip height.
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Jun ASAKAWA, Hiroyuki KOIZUMI, Shunichi KOJIMA, Masakatsu NAKANO, Nobu ...
2016 Volume 14 Issue ists30 Pages
Pa_53-Pa_59
Published: 2016
Released on J-STAGE: November 25, 2016
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A micro-solid rocket as the propulsion system for 1–10 kg-class micro-spacecraft is proposed here. The micro-solid rocket uses a boron/potassium nitrate pellet as propellant and its total impulse is about 1.5 Ns. Higher total impulse is needed for a propulsion system on small spacecraft to perform advanced space missions such as sample return, formation flight, and active debris removal. To increase the total impulse, it is necessary to increase the propellant mass. However, there is a difficulty in producing new sizes of solid propellant. The author designed a 20–50 Ns-class micro-solid rocket which uses a stack of existing multiple B/KNO3 pellets. The side of the propellant pellets was sealed with epoxy resin to prevent an abnormal combustion chamber pressure rise. As a result, all the propellant was burned without an abnormal pressure rise in all combustion tests.
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Shota ISAKARI, Shingo ONIZUKA, Yasuyuki YANO, Akira KAKAMI
2016 Volume 14 Issue ists30 Pages
Pa_61-Pa_66
Published: 2016
Released on J-STAGE: October 07, 2016
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This paper describes a solid propellant microthruster that is throttleable using laser heating. Solid propellant thrusters generally require neither tank nor valve, and accordingly have relatively high reliability due to simple structures. Nevertheless, conventional solid propellant thrusters have not been applied to attitude control or station keeping for satellites because of difficulty in throttling including start and interrupt of thrust production. Hence, we proposed to apply combustioncontrollable solid propellants, and a compact and light-weight semiconductor laser to thrusters in order to develop a throttleable solid propellant microthruster. For combustion controllable solid propellants, combustion was sustained only when external heat was supplied to burning surface. In our previous study, a prototyped 0.1 N class thruster successfully produced thrust in a vacuum, but the combustion was unstable. In this paper, to stabilize combustion, we prototyped a nozzle with a reduced target combustion chamber pressure of 0.03 MPa. Mass ratio of carbon black, which was added to absorb laser beam efficiently, was varied from 0.05 wt% to 0.5 wt%. Thrust measurement showed that the prototyped thruster successfully yields a stable thrust of 0.02 N at a laser power density of 0.98 W/mm2.
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Kazuya NAWATA, Shunya SASAKI, Tatsuya SAITO, Nobuyuki OSHIMA, Masashi ...
2016 Volume 14 Issue ists30 Pages
Pa_67-Pa_72
Published: 2016
Released on J-STAGE: November 25, 2016
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The purpose of this study is to develop a computational method to predict fuel regressions for hybrid rocket solid fuels. The shape of the flow field changes depending on the regression and vaporization of the solid fuel. This shape change and the heat flux from the combustion gas to the fuel are mutually dependent. Therefore a computational method that can accurately predicts both of the fuel regression and the heat flux is necessary to clarify the mutual dependence of them. In this study, we have developed a computational method to predict the regression phenomenon including the effect of the shape change of the flow field. The developed code predicts the regression phenomena by repeating gas-phase calculations and regression-phase calculations. The wall consisting of grids permits the flow field to be an arbitrary shape. As the first step, the complex chemical reaction was not included and numerical results were compared with a sublimation phenomenon of naphthalene in a non-combustion flow. Numerical results successfully predicted Nusselt number change due to regression qualitatively.
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Tasuku URAOKA, Yoshikazu IWAO, Yasuyuki YANO, Akira KAKAMI
2016 Volume 14 Issue ists30 Pages
Pa_73-Pa_81
Published: 2016
Released on J-STAGE: October 07, 2016
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This paper deals with a new bipropellant thruster using nitrous oxide (N2O) as an oxidizer and dimethyl ether (DME) as a fuel. Conventionally, bipropellants such as NTO/hydrazine have been used for thrusters. However, they are toxic and reactive to materials for tanks and tubes. Today, eco-friendly thrusters are required for safety and cost reduction. Hence, we proposed a new eco-friendly bipropellant thruster using N2O and DME. The bipropellant possibly reduces the cost for safety and allows easy handling because the propellant is neither toxic nor reactive to materials. N2O and DME, which are liquefied gas, are storable in a liquid form, and fed in a gaseous form simply by managing pressure and temperature. In this study, 0.4-N class thruster was prototyped to yield stable combustion and enhance performance. Using an injector consisting of a shower head and a subsequent orifice of 15 mm in diam., the prototype yielded a C* efficiency of 68.7 % and a specific impulse of 135 s at O/F=3.5.
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Hitoshi ASAKAWA, Yasuyuki YANO, Hiroaki MURATA, Akira KAKAMI
2016 Volume 14 Issue ists30 Pages
Pa_83-Pa_88
Published: 2016
Released on J-STAGE: December 09, 2016
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This paper deals with the influence of variation in combustion-chamber pressure on an arcjet plasma source for an arcjet-assisted thruster. Arcjets have been applied to chemical thrusters in order to promote combustion and augment performance. Some groups reported that combustion of solid propellant and monopropellant such as SHP163 was successfully sustained with arcjet. Arc discharge is, however, negatively influenced by variation in combustion-chamber pressure. Hence, we propose to apply active control to arcjet-assisted chemical thruster in order to stabilize combustion. The controller design requires the response of arcjets to variation in combustion-chamber pressure. In this study, we investigated the influence of step-like pressure change on arcjet exit using a combustion chamber simulator. In the simulator, pressure was suddenly increased from 0.1 to 0.35 MPa using a nitrogen-filled buffer with a burst diaphragm. At a pressure rise of 0.15 MPa with a time constant of 0.2 s, arc discharge was interrupted immediately after sudden rise in pressure of the combustion chamber simulator. From the results, arcjet was negatively affected by the combustion-chamber pressure.
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Yuichiro IDE, Takuya TAKAHASHI, Keiichiro IWAI, Katsuhiko NOZOE, Hirot ...
2016 Volume 14 Issue ists30 Pages
Pa_89-Pa_94
Published: 2016
Released on J-STAGE: November 25, 2016
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As a replacement for hydrazine, ammonium-dinitramide-based ionic liquid propellant (ADN-based ILP) has been developed by JAXA and Carlit Holdings Co., Ltd. This propellant is made by mixing three solid powers: ADN, monomethylamine nitrate, and urea. The propellant's theoretical specific impulse is 1.2 times higher than that of hydrazine, and its density is 1.5 times higher at a certain composition. Although ionic liquids were believed to be non-flammable for a long time owing to their low-volatility, recently combustible ILs have been reported. The combustion mechanism of ILs is not yet understood. The objective of this paper is to understand the combustion wave structure of ADN-based ILP. The temperature distribution of the combustion wave in a strand burner test shows a region of constant temperature. This region would indicate boiling in a gas-liquid phase. Thus, the combustion wave structure consists of liquid, gas-liquid, and gas phases. The dependence of boiling point on pressure would identify chemical substances in the gas-liquid phase. The dependence of combustion and ignition characteristics on ADN content is also discussed.
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Christian BAUER, Martin HAUSER, Oskar J. HAIDN
2016 Volume 14 Issue ists30 Pages
Pa_95-Pa_100
Published: 2016
Released on J-STAGE: October 13, 2016
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An experimental investigation has been conducted in order to evaluate whether it is possible to reduce the sensitivity of the heating effect observed in Hartmann-Sprenger Tubes by applying swirl to the nozzle flow. A reference configuration derived from literature is compared against modified cavities incorporating stems of various lengths and a setup utilizing swirl. A stem of ≈50% cavity length has been found to have a moderate effect on operating mode stability and heating effects. Subjecting a cylindrical resonator to swirling flow has demonstrated to considerably increase the amplitude of the flow oscillations, leading to detrimental effects on heating rate due to increased mass transfer from cavity to environment. The swirl configuration therefore shows potential for improving resonators used for active flow control and Hartmann-Sprenger tubes designed for larger pressure amplitudes.
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Shinji IGARASHI, Apollo B. FUKUCHI, Nobuyuki AZUMA, Keigo HATAI, Hides ...
2016 Volume 14 Issue ists30 Pages
Pa_101-Pa_105
Published: 2016
Released on J-STAGE: December 08, 2016
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We have developed a hydroxylammonium nitrate (HAN)/hydrazinium nitrate (HN)-based low-toxicity monopropellant [high-performance, no-detonation propellant (HNP)] that has safety characteristics such as no autocatalytic reaction (no autocatalytic reaction: combustion cannot continue without a source of heat) and no detonation. However, its specific impulse (Isp), a rocket engine performance indicator, was lower than that of hydrazine. Therefore, we investigated many types of compositions and found methanol to be suitable as a fuel ingredient for increasing the Isp of the developed propellant and reducing its viscosity. We produced the developed monopropellant consisting of HAN/HN/methanol/water and having a low viscosity and an Isp of 260 s at the laboratory scale.
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Akihiro IWASAKI, Kotaro MATSUMOTO, Ryosuke BAN, Shun YOSHIHAMA, Taro N ...
2016 Volume 14 Issue ists30 Pages
Pa_107-Pa_110
Published: 2016
Released on J-STAGE: October 13, 2016
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In this study, the mixing of a composite solid propellant production was investigated. Usually, propellant is mixed in batches during multi-batch processing. In this work, we demonstrated that continuous mixing with a peristaltic pump containing an artificial muscle actuator could replace batch mixing, resulting in a safe, efficient manufacturing process. In continuous mixing systems, it is important to consider the correlation between the degree of mixing of the propellant slurry and the operational variables of the pump, and we focused on the air feed pressure variable.
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Chiara BOFFA, Oskar HAIDN
2016 Volume 14 Issue ists30 Pages
Pa_111-Pa_117
Published: 2016
Released on J-STAGE: October 28, 2016
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Recent interest in hydrogen peroxide as green propellant for Space Propulsion applications justifies the efforts put in the development of numerical models and tools for the preliminary design of efficient catalysts. The multiplicity of phenomena involved, from the propellant phase change to the heterogeneous reactions, requires a step-by-step approach in order to identify key parameters and check the pertinence of the model. The phase change in porous media is the subject of the present work. In this instance the H2O2 decomposition is neglected. A multiphase mixture model for 1D, steady-state and non-reacting flows in porous media is considered. Governing equations and constitutive laws are shown and discussed. A Matlab code, relying on the SIMPLE algorithm, has been developed to solve the system of equations. Flows of oxygen-saturated water through metal foams have been considered as test cases. Results have highlighted the capabilities of the model and also the difficulties related to the solution of the gas-phase saturation equation to model the complete propellant phase change in the porous bed.
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Armin HERBERTZ
2016 Volume 14 Issue ists30 Pages
Pa_119-Pa_127
Published: 2016
Released on J-STAGE: February 08, 2017
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Steady state engine cycle analysis is commonly used in pre-design phases of liquid propellants rocket engine development. In the engine development process, a high level systems analysis, which examines the engine cycle allows a preliminary design of the engine components in terms of the operational envelope, within which the engine components are required to function. This paper compiles the general methodology and component models used in DLR's cycle analysis tools. The paper describes briefly the tool's heritage. Methods used for component modeling are described in detail. Some sample calculations of rocket engines are provided as validation examples.
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Maria P. CELANO, Simona SILVESTRI, Christoph KIRCHBERGER, Gregor SCHLI ...
2016 Volume 14 Issue ists30 Pages
Pa_129-Pa_137
Published: 2016
Released on J-STAGE: February 08, 2017
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Within the frame of a broader activity towards the use of methane as innovative propellant for rocket engines ongoing at the Institute of Space Propulsion, the efficiency of film cooling is investigated in a capacitive cooled model combustor. The combustion pressure level as well as the film applicator geometry and the film mass flow percentage are varied. The efficiency of the ambient temperature film is determined by the thermocouples installed in the copper liner along the combustion chamber axis. The fundamental dependencies from different controlling parameters are shown. Effects linked to the transient nature of the hardware are analysed. Alternative evaluation methods, which allow to deal with the transient nature of the combustion chamber are presented. Test results show the stronger influence of the blowing rate as deciding parameter for the film cooling efficiency. Increase of the blowing rate is anyhow not always connected to an improvement of the performance of the film, but a limit value could be found.
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Asato WADA, Toshiaki IIZUKA, Takahiro SHINDO, Hiroshi MAEDA, Hiroki WA ...
2016 Volume 14 Issue ists30 Pages
Pa_139-Pa_144
Published: 2016
Released on J-STAGE: February 08, 2017
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A new reaction system that uses discharge plasma of noble gas has been proposed as a substitute for the conventional solid catalyst in a reaction control system (RCS) thruster. The propellant of the thruster is a hydroxyl ammonium nitrate (HAN)-based monopropellant instead of the usual hydrazine. The proposed reaction system is a discharge plasma catalyzer (DPC) system, and a laboratory model (LM) has been developed. The DPC-LM is expected to enhance combustion via ion-molecule and radical-molecule reactions from the discharge plasma, and to enable cold-start operation. The objective of this study is to experimentally evaluate the effects of characteristics of the discharge plasma on the propellant reaction characteristics of the DPC-LM by inducing different swirl gas flow patterns, such as by varying the geometric swirl number. The plasma characteristics are evaluated in terms of the success rate of propellant reaction and the reaction delay time. The continuity of the exhaust flame was confirmed for argon gas mass flow rates from 0.125 to 0.175 g/s and an SHP163 mass flow rate of 0.3 g/s with higher geometric swirl numbers. In addition, it was found that higher geometric swirl numbers reduced the reaction delay time.
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Yuji SAITO, Tsutomu UEMATSU, Hikaru ISOCHI, Masashi WAKITA, Tsuyoshi T ...
2016 Volume 14 Issue ists30 Pages
Pa_145-Pa_151
Published: 2016
Released on J-STAGE: February 08, 2017
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A nozzle throat erosion problem occurs in developing a 15 kN-thrust class motor. To obtain a history of the nozzle throat area in a hybrid rocket static firing test, a new method is developed. Although the specific heat ratio of the combustion gas, which depends on the oxidizer to fuel ratio ξ, is necessary to calculate a nozzle throat area, it is difficult to obtain temporal ξ in hybrid rockets. A reconstruction technique, which estimates temporal ξ, needs chamber pressure, oxidizer flow rate, and nozzle throat area as input data. These two equations are solved simultaneously to acquire two convergence calculations for nozzle throat area and ξ. The new method was applied to a static firing test. The results show a typical erosion history, showing the validity of this method.
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Yasushi OHKAWA, Satomi KAWAMOTO, Teppei OKUMURA, Kentaro IKI, Yuuta HO ...
2016 Volume 14 Issue ists30 Pages
Pb_1-Pb_6
Published: 2016
Released on J-STAGE: February 13, 2016
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A flight demonstration of an electrodynamic tether (EDT) on the H-II Transfer Vehicle (HTV) is planned by JAXA. This demonstration plan is called the Konotori Integrated Tether Experiment (KITE). KITE is the first step toward the development of active debris removal (ADR) systems using EDTs. EDTs have many advantages that make them promising candidates for deorbit propulsion systems for ADR, including the absence of consumables, low electric power requirements, the absence of thrust vectoring, and easy attachment to debris. The primary objective of KITE is to demonstrate the key EDT technologies for ADR. KITE mission will be conducted prior to re-entry of the HTV-6. A 700-m-length bare tether, which is deployed from the HTV body toward the zenith, collects electrons from the ambient space plasma, and a field emission cathode on the HTV emits 10-mA-level electrons into the plasma. This collector–emitter combination can provide complete propellant-free deorbit propulsion for ADR.
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Yuriko TANIDA, Daisuke KUWAHARA, Shunjiro SHINOHARA
2016 Volume 14 Issue ists30 Pages
Pb_7-Pb_12
Published: 2016
Released on J-STAGE: July 27, 2016
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Helicon plasma sources have been studied in many fields across science and technology because they can supply high-density plasmas with a broad range of external operating parameters. In our laboratory, we aim to develop a completely electrodeless electric thruster, which is expected to have a high efficiency and a long lifetime, leading to be useful on a deep space exploration. In order to demonstrate and optimize this thruster system, it is important to have detailed distributions of plasma flow. Laser induced fluorescence (LIF) diagnostics, which can measure velocity distribution functions of particles, has advantages from a view point of, e.g., high resolution in time and space, and it can determine an absolute particle velocity and its temperature in addition to its relative density. Here, we have been developing a LIF measurement system using a Multi-Pixel Photon Counter (MPPC) for a multi-channel system. Argon ion velocity depending on the magnetic field gradient in a downstream region is measured by this LIF system.
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Hiroyuki KOIZUMI, Hiroki KAWAHARA, Kazuya YAGINUMA, Jun ASAKAWA, Yuich ...
2016 Volume 14 Issue ists30 Pages
Pb_13-Pb_22
Published: 2016
Released on J-STAGE: July 29, 2016
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Initial flight operations of the miniature propulsion system I-COUPS (Ion Thruster and COld-gas Thruster Unified Propulsion System) are presented with problems found in space and its countermeasures for them. The I-COUPS was developed by the University of Tokyo and installed on a 70 kg space probe, PROCYON as main propulsion system to verify propulsive capability of the first micropropulsion in deep space. The PROCYON was successfully launched on December 3rd, 2014 and inserted into an orbit around the Sun. The PROYON project team started flight operation on the interplanetary orbit. Up to today, the cold-gas thrusters have successfully conducted unloading maneuvers since the launch. The ion thruster overcame several problems and achieved 223 hours operation with the averaged thrust of 346 μN. The I-COUPS will become the first electric propulsion and reaction control system operated on a small space probe (<100 kg) on an interplanetary orbit.
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Chris VOLKMAR, Ubbo RICKLEFS, Peter J. KLAR
2016 Volume 14 Issue ists30 Pages
Pb_23-Pb_32
Published: 2016
Released on J-STAGE: July 27, 2016
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In this paper, we show a numerical model of gridded radio-frequency (RF) ion thrusters. The model consists of a set of self-consistently coupled equations based on conservation of charge, energy, and mass inside the system. Those 0D models are again coupled to a 3D quasi-stationary electromagnetic field solver which offers the opportunity of evaluating arbitrary induction coil and discharge chamber geometries in fairly reasonable simulation time. Several input parameter sets can be computed in parallel due to multi-core implementation. Therefore, the model presented can be regarded as a toolbox for ion thruster engineering purposes and rapid virtual prototyping. The model predicts electrical parameters such as thruster and plasma impedance as well as propulsive performance data. It is thus possible to use it for finding optimized coil and chamber geometries together with optimized input parameters (coil current, volumetric propellant flow rate, and extraction grid voltages) in order to obtain improved mass and electrical efficiency. To prove the validity of the model, performance mappings experimentally performed on a RIM-4 RF ion thruster assembled at the University of Giessen are used to verify the computed data.
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Jens SIMON, Uwe PROBST, Peter J. KLAR
2016 Volume 14 Issue ists30 Pages
Pb_33-Pb_39
Published: 2016
Released on J-STAGE: July 27, 2016
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To ensure the successful operation of a thruster using inductively-coupled plasmas (e.g. radio-frequency ion thrusters (RIT)), a high efficiency and high performance radio-frequency (RF) power supply is crucial. For this purpose, the supply needs to guarantee highly efficient RF signal generation and transfer to the thruster coil and furthermore an optimal power coupling between coil and plasma. In this paper we propose a high efficiency, high performance approach of generating RF power by using a resonant converter design. Due to a compensation of the load's inductance, resonant switching behavior becomes possible and thus, switching losses can be significantly reduced. The drive signals of the power semiconductors are generated by a switching frequency-adapting, load-controlled algorithm, which keeps up a quasi-resonant state. Due to high-speed tracking of the resonance frequency and phase, RF generation can be adjusted to altered load conditions within a few RF-cycles. A digital FPGA-based implementation guarantees precise period determination and control signal generation, since the programmed control-algorithm is executed in hardware. The proper functioning of the developed radio-frequency generator (RFG) concept is verified through performance mappings, recorded when supplying a RIM-4 RF ion thruster developed at the University of Giessen. In the long run, this RFG concept shall be employed for driving RIT more efficiently.
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Yoshinori TAKAO, Hiroyuki KOIZUMI, Yusuke KASAGI, Kimiya KOMURASAKI
2016 Volume 14 Issue ists30 Pages
Pb_41-Pb_46
Published: 2016
Released on J-STAGE: September 15, 2016
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To investigate electron extraction through the orifices of a microwave discharge neutralizer, three-dimensional particle simulations have been conducted. The numerical model is composed of a particle-in-cell simulation with a Monte Carlo collision algorithm for the kinetics of charged particles, a finite-difference time-domain method for the electromagnetic fields of 4.2-GHz microwaves, and a finite element analysis for the magnetostatic fields of permanent magnets. The distribution of the current density on the orifice plate obtained from the numerical model is in a reasonable agreement with the measurement result in an experiment. Moreover, the numerical results have indicated that the electrostatic field of the plasma has a dominant influence on the electron extraction although the electrostatic field produces the opposite force of extraction from the bulk plasma toward the orifice plate. The combination of the sheath potential barrier and the magnetostatic field yields the electron trajectories of extraction.
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Eduardo FERNANDEZ, Caleb DOWDY, Jacob ALEY
2016 Volume 14 Issue ists30 Pages
Pb_47-Pb_55
Published: 2016
Released on J-STAGE: September 15, 2016
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A two-dimensional model of the Hall thruster with kinetic, non-magnetized ions, and fluid electrons is presented. The model dynamically evolves azimuthal flows and fluctuations, differing from standard hybrid models that assume axisymmetry and resolve quantities in the radial and axial coordinates only. Unlike those descriptions, which typically use adhoc cross-field electron transport parameters in order to sustain the discharge, the present model relies on classical transport and fluctuations generated within the plasma. A number of low-frequency wave modes are captured in the simulation, from the lowest (few kHz) “breathing mode”, to waves on the order of a few hundred kHz. At issue is the characterization of the various modes with regards to their instability mechanisms, their spectral signatures, their dependence on plasma inhomogeneity along the channel, and their role in cross-field electron transport. Simulations show that gradient-driven drift instabilities emerge downstream of the peak of the magnetic field, as predicted by linear stability analysis, while strong, ionization driven fluctuations take place upstream. While fluctuations result in fluctuation-driven transport, they do not drive sufficient current to match experimental measurements. Electron transport is reduced in the strong magnetic field region to near classical levels, in qualitative agreement with experiments.
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Kenta HIRAMOTO, Yoshinori TAKAO
2016 Volume 14 Issue ists30 Pages
Pb_57-Pb_62
Published: 2016
Released on J-STAGE: September 17, 2016
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We have investigated ion beam extraction mechanism using two dimensional axisymmetric particle-in-cell simulations with Monte Carlo collisions (PIC/MCC) under various conditions of grid structures. The calculations are carried out for both the plasma region and the vacuum region simultaneously, where the former is 5.0 mm in radius and 10 mm in length and the latter is 6.0 mm in radius and 20 mm in length. The PIC/MCC results have shown that the ion beam current is affected by the gap distance of the grids and thickness of the screen grid, but is not affected by the hole diameter of the screen grid. Moreover, the optimum value of the difference between the hole diameter of the screen grid and that of the accelerator grid is 1.4 mm at various grid gap distances and thicknesses under the following conditions: the gas pressure is 3.2 mTorr, the absorbed power is 500 mW and the beam voltage is 1100 V.
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Kaito NAKAGAWA, Yoshinori TAKAO
2016 Volume 14 Issue ists30 Pages
Pb_63-Pb_68
Published: 2016
Released on J-STAGE: September 17, 2016
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We have investigated the plasma production in a micro RF ion thruster, where the plasma source is 5.0 mm in radius and 10 mm in length. To find the optimum condition to generate the RF plasma in the wide range of frequency f = 1-1000 MHz and pressure p = 0.1-10.0 Pa, we employ the equivalent circuit model, where the global model is also incorporated to obtain the plasma parameters self-consistently. The numerical results have indicated that capacitive coupling dominated over inductive coupling and the sheath resistance has a significant influence on the power coupling efficiency at lower frequency. The power coupling efficiency could be more than 80% at higher frequencies (> 30 MHz).
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Stefan SCHARRING, Raoul-Amadeus LORBEER, Hans-Albert ECKEL
2016 Volume 14 Issue ists30 Pages
Pb_69-Pb_75
Published: 2016
Released on J-STAGE: September 30, 2016
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Aiming for the generation of high-precision thrust in the µN range, focused high-intensity laser pulses are used employing the recoil of the jet of the ablated material. Whereas a single laser pulse yields an extremely low impulse bit in the range of several nNs, a broad thrust range can be accessed by the variation of the laser pulse repetition rate up to several hundreds of kilohertz. A detailed laser parameter study is carried out for aluminum and gold as propellant varying the pulse length from 100 fs to 10 ns and the fluence from 0.09 to 23.8 J/cm2. Two different regimes of thruster operation with respect to laser pulse length and specific impulse are identified. Irrespective of the pulse length regime, optimum impulse coupling is found at laser spot fluences around 2 J/cm2 for aluminum and 20 J/cm2 for gold, respectively, with coupling coefficients in the range of 25 to 40 μN/W. For ultrashort pulses, jet velocities are rather small yielding a specific impulse in the range of 70 s to 200 s, whereas for longer pulses beyond ≈ 100 ps, Isp is found to be in the range of 500 to 1000 s and beyond enabling low propellant consumption. However, ultrashort-pulse laser ablation might be favorable since material can be removed very smoothly which might contribute to very low thrust noise.
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Hiroki WATANABE, Takanori DEGUCHI, Shuka TAKEDA, Yuki MIURA, Masanori ...
2016 Volume 14 Issue ists30 Pages
Pb_77-Pb_82
Published: 2016
Released on J-STAGE: September 30, 2016
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To liberate Hall thrusters from the drawbacks associated with dispenser hollow cathodes, we construct and experimentally evaluate an outer-coil-type radio frequency (RF) plasma cathode and an inner-coil-type RF plasma cathode. The influence of the coil configuration on the electron-emission characteristics of the RF plasma cathodes is significant. Compared to the inner-coil-type RF plasma cathode, the outer-coil-type RF plasma cathode achieved higher electron- emission performance. For the outer-coil-type RF plasma cathode, we obtained an anode current of 3.3 A at an RF power of 140 W, a xenon mass flow rate of 0.3 mg/s, and an anode voltage of 58 V. The anode current is sufficiently high to operate a 1-kW class Hall thruster. The gas utilization factor for the outer-coil-type RF plasma cathode is comparable to that for a conventional dispenser hollow cathode. On the other hand, the electron production cost for the outer-coil-type RF plasma cathode is four times higher than that for the hollow cathode. Thus, there is a need to improve the power consumption for application of RF plasma cathodes to Hall thrusters.
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Kento HOSHI, Hirotsugu KOJIMA, Hiroshi YAMAKAWA
2016 Volume 14 Issue ists30 Pages
Pb_83-Pb_89
Published: 2016
Released on J-STAGE: November 26, 2016
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We perform the first three-dimensional particle-in-cell simulations of the electric solar wind sail, which is a recently proposed new propulsion system. We investigate the potential structure around the tethers of the electric solar wind sail. As a result, the potential distribution is greatly lower than the potential expression proposed by the previous studies. We proposed the new method to estimate the potential around the tether numerically without performing the time-consuming PIC calculation. The proposed numerical solution has a very good agreement with PIC results under two difference plasma environments, and its agreement shows the validity of the method. The proposed method is useful to estimate the thrust of E-sail when a precise thrust operation is required.
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Kateryna AHEIEVA, Kazuhiro TOYODA, Mengu CHO
2016 Volume 14 Issue ists30 Pages
Pb_91-Pb_97
Published: 2016
Released on J-STAGE: November 26, 2016
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This paper will describe the development of the Vacuum Arc thruster (VAT) for micro and nano satellites with the main purposes of: attitude control, maintaining orbit-satellite deorbiting. Firstly, this VAT thruster is integrated on-board student microsatellite Horyu-IV, which was developed at the Kyushu Institute of Technology (KIT), Japan. The satellite will be launched in the fiscal year 2015 by H2A rocket. This paper describes the principles of VAT as a direct drive from High Voltage Solar Array (HVSA). Expected performance of this vacuum arc thruster with passive ignition (space plasma condition) was measured. We propose a method for improving the thruster efficiency. Results show that the impulse bit was of the μNs order, and the thrust–56 nN– and efficiency– 2.5%–were calculated. Moreover, it was found that the impulse bit changes with the applied voltage. A new CFRP material was developed and used as a cathode for the purpose of improving efficiency. Discharge characteristics of current and the arc rate are presented. Measurements of impulse bit were also done for the VAT in configuration with two different propellants (commercial and new CFRP) and with a permanent magnet of 300 mT.
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Kaoru KAKINUMA, Masafumi FUKUNARI, Toshikazu YAMAGUCHI, Yusuke NAKAMUR ...
2016 Volume 14 Issue ists30 Pages
Pb_99-Pb_103
Published: 2016
Released on J-STAGE: December 09, 2016
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This paper proposes a two-stage-to-orbit launch system comprised of a Microwave Rocket 1st stage and a Microwave Thermal Rocket 2nd stage. The air-breathing 1st stage improves payload fraction relative to a single-stage-to-orbit system and carries the 2nd stage above the atmosphere and into range of its beam director. For the 1st stage task, a Microwave Rocket is superior to an unmanned aerial vehicle because it is simpler, faster, and reaches higher altitude at higher speed. In addition, we present a new trajectory that eliminates power beaming at low elevation angles and improves system performance. This combination of factors reduces the propellant needed in the 2nd stage, which in turn increases payload fraction by a remarkable factor of 3 times.
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Hiromasa TAKENO, Yusuke TOGO, Tomohiro KATSURA, Yasuyoshi YASAKA, Kazu ...
2016 Volume 14 Issue ists30 Pages
Pb_105-Pb_109
Published: 2016
Released on J-STAGE: September 30, 2016
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Traveling wave direct energy converter (TWDEC) was proposed as an efficient energy recovery device for fast protons produced by D-3He fusion reaction. The application of TWDEC to fusion propulsion system was also studied, and its significant subject was miniaturization of the device. In TWDEC, there is a trade-off between device size and efficiency, and an employment of a decelerator of constant deceleration scheme is promising to realize miniaturization. The paper experimentally examines one of the working characteristics of the constant deceleration scheme by using a relative phase control method called active decelerator. The results of the experiment and corresponding numerical orbit calculation are consistent with the theory of the constant deceleration scheme.
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Atsushi YAMAGUCHI, Atsushi KIBE, Naoji YAMAMOTO, Taichi MORITA, Hideki ...
2016 Volume 14 Issue ists30 Pages
Pb_111-Pb_116
Published: 2016
Released on J-STAGE: November 26, 2016
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A cavity ring-down spectroscopy method to measure sputter erosion and estimate the lifetime of electric propulsion systems was developed. We measured sputtered aluminum atoms from the acceleration grid of an ion thruster. An external-cavity diode laser was used as the probe laser to measure the aluminum transition line from ground state to upper state, at 394.512 nm (vacuum). The erosion rate of the acceleration grid was estimated to be 259 ± 99 ng/s.
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