Engine system analysis upon a Rocket-Based Combined Cycle (RBCC) engine for the booster stage of a TSTO launch vehicle was conducted, to evaluate the effects of hydrocarbon fuel type upon the performance of the engine and the whole launch vehicle. A higher thrust and cooling performance with methane was shown, while some technical challenges were also noted. Combination of the scramjet flow pass and the rocket engines, i.e., rocket engine being embedded into the scramjet flow pass or separately mounted to the airframe, was varied to show that the separated configuration mitigated cooling problem, and resulted in a better performance for mid-speed regime.

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A system that integrates on hydrogen-base all of energy on propulsion and power needed in a space transportation system is one of mechanisms for operating reusable space transportation system frequently. In this paper, the concept design for the integrated system is presented to discuss the effectiveness of the system. As evaluation functions, (1) weight, (2) cost, and (3) turn-around time (operability) are considered. Modeling of the energy system is conducted focusing energy flow and connections between elements that configure the energy system. Assuming a space transportation mission and regarding the required energy function in the mission as boundary conditions, conceptual designs of the integrated energy system using the developed model are performed. There is a trade-off relationship between the three evaluation functions. For weight reduction, integration of fuel species and the scale of the energy storing elements, those are an accumulator and a battery, is effective. For reduction of the turn-around time, toxic fuels have to be removed. An adoption of batteries reduces the number of connections between the elements so that it contributes to improvement in operability. As for cost, the part related to the manufacturing depends on weight and that related to the operation does on the turn-around time in the present cost model. Hydrogen and oxygen-based integrated energy system, although additional elements are to be necessary and a disadvantage in terms of weight is caused, is effective in reducing the operating costs and the turn-around time in case that a large amount of high frequency transport is desired because it is possible to improve the operability significantly.

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Bioethanol is a suitable candidate as a reusable rocket engine fuel because it is carbon neutral and environmentally friendly. Based on the JAXA future transportation reference system, the feasibility of a gas generator cycle fueled by bioethanol was examined in this paper. The result is considered to be a general design guideline of an ethanol-fueled rocket engine. A combustion pressure of 7 MPa, an expansion ratio of 25, a turbine efficiency of 0.6 and a pump efficiency of 0.7 enabled the engine to satisfy the required specifications with an Isp efficiency of 0.96. In addition, a flow experiment of bioethanol under high temperature and high pressure was conducted. Resulting from the EPMA analysis, the presence of sulfur and coking was recognized on the copper-alloy tube.

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A theoretical model for the ground reflection from a correlated, extended source and including atmospheric turbulence [K. L. Gee et al., Proc. Mtgs. Acoust. 22, 040001 (2014)] is used to correct spectra measured over snow-covered terrain during a horizontal solid rocket motor firing. A sensitivity analysis reveals that at moderate distances, the relative sound pressure level changes are more sensitive to ground effective flow resistivity and microphone height, whereas at long range, the turbulence parameters and ground impedance have greater impact.

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Near-field characterization of the acoustical environment near rockets has often involved extrapolating far-field measurements. However, because far-field amplitude data reveals only limited information about source characteristics, a vector intensity measurement system and analysis package has been developed to examine source features more directly. This paper describes the development of the measurement and analysis capability and its application to a horizontal firing of a GEM-60 solid propellant rocket motor firing conducted at ATK Space Systems near Promontory, Utah. An analysis of near-field intensity data provides insight both into the spatial extent and principal radiation lobe as a function of frequency. For 50 Hz, the far-field spectral peak frequency in the maximum radiation direction, the dominant source region derived from tracing the near-field intensity vectors spans 17-32 nozzle diameters, with peak radiation at ~68°. At high frequencies, the radiation results from a more contracted region that occurs farther upstream and is directed at about ~85°. These results point to the potential utility of near-field vector intensity measurements, in part because the near-field environments represented do not agree with historical far-field data-based models.

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Development of the next-generation space flight vehicles has prompted a renewed focus on rocket sound source characterization and near-field propagation modeling. Improved measurements of the sound near the rocket plume are critical for direct determination of the acoustical environment both in the near and far-fields. These measurements are also crucial inputs to empirical models and utilized in validating computational aeroacoustics models. NASA's SP-8072 acoustic load prediction model [K. M. Eldred, NASA SP-8072 (1971)] is a widely used empirical model for predicting liftoff acoustics. The model implements a Distributed Source Method, which predicts the loading as the sum of the radiated acoustic pressure field from each source distributed along the plume. In this paper, the resultant sound distributions are compared among the original core length definition in SP-8072, a modification proposed by Varnier [J. Varnier, AIAA Journal, Vol. 39, No. 10, (2001)], and four full-scale solid rocket motor tests with measurements covering a spatial region of 9 to 56 nozzle diameters downstream of the nozzle. The results of the comparisons support the original Eldred core length definition on an overall level basis. However, analysis of the measured spectra unveil shortcomings within both definitions.

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Effects of solid particles in the gas-particle multiphase compressible turbulent mixing layer were investigated by direct numerical simulation, compared with the single phase turbulent mixing layer of which the net density ratio (sum of gas and particle densities) was set to be the same as the multiphase case. Here, the situation of solid rocket motor was considered, so that the Mach number and the density (temperature) ratio were set to be high and far from unity, respectively, and the solid particles were distributed only in one side of the mixing layer for the initial condition. Three-dimensional compressible gas-particle-multiphase Navier-Stokes equations in the Euler-Euler formulation were solved with an alternative weighted essentially non-oscillatory scheme with a positivity preserving limiter of the particle density for the multiphase flow computation. It was found that the existence of solid particles leads to the sparse structures of turbulence and suppresses the fine scale turbulent structures, especially for the JET side, which corresponds to the side where particles were initially distributed and the gas density was low. The growth rate of the mixing layer thickness for multiphase flow became smaller due to the change in flow. The change in flow structures also affected the properties of acoustic waves. The Mach wave like structures do not appear for the initially particle existing side for the multiphase flow due to the suppression of fine scale turbulent structures, compared with single phase flow.

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This paper demonstrates the applicability of serviceability function to the studies of satellite constellation tradeoffs for detecting damaged area distribution in the emergency response after disaster eruption. The example studies indicate that constellations of synthetic aperture radar (SAR) satellites with different local sun time orbits can be effective for early response following a disaster. This study also proposes the introduction of visibility and interpretability criteria for disaster monitoring, beginning with the addition of necessary criteria for use in the emergency response phase. Here, image-based products that combine the change detection by SAR and compensational use of optical imagery, social database and matured operators are proposed as possible candidates for employment during this phase, which are to be evaluated by the criteria.

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Landsat-8 was launched on February, 2013, and the data is freely available and it is used in various fields of application at present. In November 2013, an eruption created a new island near Nishinoshima. The new island grew rapidly and ended up joined with Nishinoshima. This volcanic activity has been observed by Landsat-8 every 16 days except on cloudy days. Meanwhile, the FORMOSAT-2 of the Taiwanese National Space Program Office (NSPO) was launched on May, 2004 and is currently operational. It collects high-resolution panchromatic and multispectral images during its sun-synchronous circular orbit. Japan and environs is located between its orbital tracks. As a result most of the images taken of Japan are observed from an angle of over 30°. It is possible to supplement images in the gap period of Landsat-8 observation with those from the FORMOSAT-2.

This study deals with the monitoring of Nishinoshima volcanic activity using Landsat-8 time series images and additional information obtained from FORMOSAT-2 images is also discussed.

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This paper describes an analysis result for the extraction and elevation measurement of the gravel bar as micro-topographical feature by ALOS data for the purpose of the restoration of the past water area in the Old Juyan Lake located in the lower reaches of the Heihe of northern China. Extraction and elevation measurement of about 20 gravel bars distributed from 900 m by 926 m in study area were carried out by newly created DSM and ALOS pan-sharpened image. The verification of validity of the elevation and age identification of gravel bars was performed by field survey measurements. Moreover, presumption of the Old Juyan Lake and reconstruction of their change for past about 7400 years were carried out. Finally, the influence of climate change and human activities to the change was discussed.

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To provide user-friendly, free, and open moderate-resolution earth observation data, namely, Landsat-8 data, a new graphical user interface (GUI), LandBrowser, which disseminates data using Open Geospatial Consortium (OGC) standards, such as Web Coverage Service (WCS) and Web Map Service (WMS), was launched on December 4, 2014. LandBrowser is used to search and save images. It does not require previous knowledge of image processing, image sensors, date of observation, or geography, and can share information interactively. Statistics, such as the number of downloaded image data, are introduced. To improve position accuracy, Ground Control Points (GCPs) are measured and compared with obtained data. Position accuracy is indispensable to change detection. GCPs were selected in Tsukuba, Hokkaido, and Tohoku areas. Most of the errors were within one pixel. As an example of high-level data application, another GUI called Hotarea, which detects hot spot activities, such as volcanoes, forest fires, and factories, was created. Facebook and Twitter were used to introduce events and activities, methods of use of our disseminated images, and interesting images of the day or related topics. Facebook articles were updated daily. Our future plan to improve user interfaces, which will treat UNIFORM-1 data and CIRC data of Daichi-2, is introduced.

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In the U.S., conflicts are occurring between indigenous people and research facilities for space science. To solve and prevent such conflicts, it is necessary to fill the gaps of concepts over various issues between indigenous people and researchers. Also, increasing the number of indigenous scientists who can serve as moderators may improve the situation.

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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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, I_sp 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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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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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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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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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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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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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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