Authorship：Lead author,　Corresponding author   Language：English   Publishing type：Research paper (scientific journal)

<p>Epsilon-4, which was the Enhanced Epsilon launch vehicle optional configuration with a multi-satellite mount structure, was successfully launched with seven satellites in January 2019. The seven satellites consist of 200 kg-class small-satellite, three micro-satellites and three CubeSats in the Innovative Satellite Technology Demonstration-1 program. The post flight analysis was conducted. The chamber pressure and thrust of three main motors and chamber pressure of SMSJ and SPM were analyzed to confirm the effectiveness of design and production methodologies. All solid propulsion systems for Epsilon-4 showed a very good behavior during the flight. As a result, the validity of the design and operation of the Enhanced Epsilon launch vehicle was confirmed.</p>

DOI： 10.2322/tastj.19.400

CiNii Article

CiNii Research

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Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.2514/1.A34695

DOI： 10.2514/1.A34695

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A numerical set up has been developed to predict the axial distribution and time history of fuel regression rates with the aim of comparing numerical results with the experimental data. The numerical approach is based on a quasi-one-dimensional CFD strategy, which is coupled with chemical equilibrium combustion; the wax-fuel regression rate has been simulated with the liquefying fuel theory developed by Karabeyoglu. In this model, local radiation heat transfer from both the hot gas combustion molecules (CO, CO2, and H2O) and soot particles is considered. It is shown that, when radiation heating of the fuel is taken into account, the calculated fuel regression rates exhibit values and axial profiles close to the ones observed experimentally.

DOI： 10.2514/6.2020-3768

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For the purpose of constructing a reliable and reasonable evaluation method of the safety of hybrid rockets, the evaluation of the safety distance for blast wave is conducted. Physical phenomena leading to the blast of hybrid rocket propellants are extracted and modeled recognizing fuel fragmentation and dust explosion as key phenomena which cause blast waves. In order to evaluate the amount of fuel dust, of the particle size less than or equal to 500 μm, a particular correlation is used. The correlation between dust mass fraction and the ratio of the applied energy to the absorbed energy by the fuel is represented by utilizing existing experimental data. Also indefinite parameters, four energy efficiencies during the explosion processes, are identified by reproduction of existing experimental data of blast pressure measurement tests. It is found that, by using the adjusted model parameters, the present model can reproduce the existing safety distance data of hybrid rocket propellant blast for various impact velocities. The correlation about dust mass is valid for various fuels and, with this, it is also implied that the evaluation model for blast is applicable to various situations.

DOI： 10.2514/6.2019-3917

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Authorship：Lead author,　Corresponding author   Language：English   Publishing type：Research paper (scientific journal)

<p>The development of enhanced propulsion system for the next Epsilon rocket was progressed. The development of Enhanced Epsilon is mainly the renewal of the second stage, and also includes each subsystem's improvement. The second stage motor M-35 was newly designed and manufactured. In order to verify the design, the static firing test of the second motor M-35 under the condition of vacuum ambient was conducted in 2015. The JAXA successfully launched the first Enhanced Epsilon launch vehicle. All solid propulsion systems for the Enhanced Epsilon launch vehicle showed a very good behavior during the flight</p>

DOI： 10.2322/tastj.17.289

CiNii Article

CiNii Research

Other Link： https://ci.nii.ac.jp/naid/130007642475

Language：English   Publishing type：Research paper (scientific journal)

A breadboard model (BBM) of an altering-intensity swirling-flow-type (A-SOFT) hybrid rocket engine was recently developed, and static firing experiments of the BBM were performed under various axial and tangential oxidizer mass-flow rates. A-SOFTs are intended to control thrust at an optimal oxidizer-to-fuel-mass ratio and achieve a high baseline of regression rates with a fixed motor configuration. This is possible by controlling both the oxidizer mass-flow rate and the effective geometric swirl number. A simple model based on a continuous and monotonic function of oxidizer mass flux and effective geometric swirl number was able to accurately predict the performance of experiments conducted on the A-SOFT BBM. The local fuel regression behavior in the axial direction of the A-SOFT BBM was shown to be like that of a swirling-oxidizer-flow-type hybrid rocket engine. Combustion efficiency was evaluated indirectly using equations for the efficiencies of thrust and specific impulse to eliminate errors due to local pressure shifts and the centrifugal force in swirling flows. In most cases, this indirect method compensated for the overestimations of c efficiency that resulted by directly using chamber-pressure data.

DOI： 10.2514/1.B36889

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Authorship：Lead author   Language：English   Publishing type：Research paper (international conference proceedings)

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

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<p>Altering-intensity Swirling-Oxidizer-Flow-Type (A-SOFT) hybrid rocket engine (HRE) was proposed as a technique to solve problems of current hybrid rockets. It uses axial and tangential oxidizer injections and their mass flow rates are manipulated independently to control the thrust and O/F. The visualization experiment of combustion of gaseous oxygen (GOX) and polymethyl methacrylate (PMMA) of A-SOFT HRE is carried out under combustion pressure of 1 bar. Combustion gas flow in the combustor is captured by high speed cameras whose fps is 30000. In order to capture the nature of the flow field quantitatively, obtaining velocity profiles by image analysis applied to its visualization images is effective. Basic image analysis method is Direct Cross Correlation method and, for high precision and spatial resolution, Correlation Based Correction is applied recursively. Averaged velocity profiles are obtained by averaging the calculation results of 10000 images corresponding to the actual time duration of 0.3 s. Flames with strong white light emission are characteristic of hybrid rockets and regarded as traceable markers for the image analysis. These luminous flames are considered to flow in the boundary layer and velocity of flames in the boundary layer is important to capture the characteristics of combustion flow field. Axial velocity of the luminous flame is proportional to the total mixed mass flow rate for each location x. Tangential velocity is proportional to angular momentum given by tangential GOX injection and inversely proportional to the total mixed mass flow rate for each location x. And how much GOX injected in axial and tangential directions are mixed is speculated by velocity profiles. It is found that mixing of GOX injected in axial and tangential directions occurs at almost the same constant rate regardless of the ratio of GOX mass flow rate injected in axial and tangential directions.</p>

DOI： 10.1299/jfst.2018jfst0029

Scopus

CiNii Article

CiNii Research

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DOI： 10.1061/(ASCE)AS.1943-5525.0000778

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Authorship：Lead author   Language：English   Publishing type：Research paper (international conference proceedings)

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Burning tests were carried out for the 5-kN-Thrust swirling-oxidizer-flow-Type hybrid rocket engine to establish the engine technology and to obtain the fuel regression characteristics. The engine attained stable burning without any combustion oscillation and C efficiencies of more than 98% under the 5-kN-Thrust conditions. The maximum timeaveraged thrust realized was 4.4 kN, which was lower than 5 kN. This was attributed to the regression rate being lower than the value estimated by regression rate correlation. The fuel regression behavior of this engine was characterized as the flow-development region near the grain leading edge and the fully developed turbulent flow region. The axial length of the flow-development region varied from that observed in the laboratory-scale engine, and it depended on the oxidizer mass flow rate, grain initial port diameter, and swirl strength. The controlling parameter of the regression rate in the fully developed flow region was modified to correlate with the grain port diameter. By effecting these modifications, the accuracy of the regression rate prediction was improved. The range of the applicable controlling parameter of the modified regression rate correlations was extended for use for larger-scale engines.

DOI： 10.2514/1.B36239

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Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1504/IJAL.2016.074912

Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1299/jfst.2016jfst0003

CiNii Article

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Language：English   Publishing type：Research paper (international conference proceedings)

A bread board model of Altering-number Swirling Flow Type (A-SOFT) hybrid rocket engine has been newly developed and the static burning tests have been conducted. A-SOFT hybrid rocket engines have controllability of both thrust and O/F without any replacement of components in the engine. This advantage is acquired by controlling oxidizer mass flow rate and effective swirl number. The purpose of this set of experiments is to confirm the continuity, monotonousness and predictability of the performances of A-SOFTs. The A-SOFT BBM showed a favorable fuel regression behavior. The fuel regression averaged along spatial direction fit the shape of regression rate function proposed before the experiments within ±3.5% errors, and the fuel regression rates are continuous and monotonous along swirl number and oxidizer mass flux. The O/F and thrust data also respectively fit the prediction formulas within ±4.2% and ±4.7% errors. c* efficiency is evaluated with the Isp efficiency - nozzle efficiency ratio ηIsp/ ηCF in order to compensate the pressure sensing errors caused by the centrifugal forces of swirling flows. Though ηIsp/ ηCF in the cases of weak swirl injection was clearly larger than in the cases of axial injection, its dependence on effective geometric swirl number was not clear in strong swirl conditions.

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Regarding center-perforated composite solid propellant grains, the radial linear burning rate of the propellant often depends on its location in the web. In most cases, the burning rate at the middle of the web is highest along the radial direction. This deviation of the linear burning rate along the radial direction is called the midweb anomaly, the hump effect, and so on. The phenomenon is the result of multiple factors (e.g., orientation of oxidizer particles, existence of binder layers, and the effect of those factors on combustion and heat conduction in the solid phase). The physical cause of this phenomenon has not been understood in sufficient detail. In this study, the effect of oxidizer particle orientation on the burning rate is studied. The particle orientation in inert solid propellant grains is observed by micro-focus X-ray computerized tomography. To simplify the indefinite shape of actual ammonium perchlorate particles, simple cylindrical inert particles were mixed into inert propellant slurry. As a result the particle orientations were observed in center-perforated and solid cylindrical inert grains. The orientation of the particles seemed to be along the isochronous surface, which is formed during the casting process. One physical correlation between the particle orientation and the mean or local burning rate is investigated by numerical simulation from the point of view of the combustion surface configuration, which is formed by the different burning rates of the propellant ingredients. The results of the simulation suggest that the orientation of oxidizer particles in line with the burning direction increase the mean burning rate by forming and developing uneven surfaces.

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2016014195

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DOI： 10.1615/IntJEnergeticMaterialsChemProp.2017011994

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Authorship：Lead author   Language：English   Publishing type：Research paper (international conference proceedings)

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Language：English   Publishing type：Research paper (international conference proceedings)

DOI： 10.2514/6.2016-0583

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Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.20910/IJASE.2015.2.2.029

Language：English   Publishing type：Research paper (international conference proceedings)

A multi-objective genetic algorithm (MOGA) has been applied to the multidisciplinary design optimization (MDO) of a launch vehicle (LV) with a hybrid rocket engine (HRE) to investigate the ability of an HRE to serve as a sounding rocket from various perspectives. In this study, the flight evaluation was enhanced to 3-degree-of-freedom (3DoF) in order to consider the equations of motion for horizontal and vertical motion and rotation of the LV. In the consideration of the rotation of the LV, the time variation of the center of gravity due to the fuel burn was estimated. The non-dominated sorting genetic algorithm-II (NSGA-II) was used to solve multi-objective problems (MoPs). Four design problems were examined in order to understand the practical physics of hybrid rocket. As a result, tradeoff information was observed for all design problems. The results for the present four design problems indicate that economical performance of LV is limited with the HRE in terms of the maximum altitude and maximum downrange distances achievable. The hypervolume, which was used as the metric to evaluate the difficulty of the design problems, reveals that the convergence of the solutions for not only altitude maximization in the case of a vertical launch but also the maximization of downrange at higher target altitudes was affected by the severe limitation. To observe the dependence of the design problems on the constraint, the design problems were visualized using a colored parallel coordinates plot (PCP), and the LV geometries determined from the nondominated solutions were successfully examined.

DOI： 10.1109/CEC.2015.7256948

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Language：English   Publishing type：Research paper (international conference proceedings)

Conceptual design of a launch vehicle with a hybrid rocket engine (HRE) has been implemented using design informatics approach in order to investigate the feasibility of a single-stage hybrid rocket. Two test design problems were formulated by using two objective functions: maximization of downrange and minimization of initial gross weight, seven design variables which describe geometry and initial conditions, and one constraint relative to target altitude. The optimization result reveals the economical performance of hybrid rocket is limited with HRE in terms of the maximum downrange achievable. Moreover, the data-mining result indicates the mechanism of design-variable behavior.

DOI： 10.1515/tjj-2015-0038

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (scientific journal)

The flow field of multi-parallelized swirling flow inside a circular pipe was investigated numerically. Two types of swirling flow configuration are considered. One type is the co-rotating type. Four co-rotating swirls are arranged at the vertex position of square in this type. The other type is the counter-rotating type which consists of two pairs of swirls having opposite swirl rotations. Each pair is arranged diagonally at the vertex position of a square. By coupling the discrete vortex method and boundary element method, unsteady flow simulation is performed. Swirl modeling with vortex elements is used in this simulation and its validity is confirmed. From the simulation results, in the co-rotating type, the four swirls interact and their shape is deformed. Each vortex motion vanishes rapidly in the downstream region. Finally, they are turned into a single swirling flow. In counter-rotating type, each vortex motion is maintained a little bit longer than co-rotating type, and their shape is not so deformed. However, the flow patterns are changed completely in the downstream region. The swirling velocity of each swirl mostly vanishes. Finally, they are turned into an axial flow. For the investigation of the mixing promoting effect due to parallelizing swirls, particle tracking simulations are performed in the co-rotating type and the counter-rotating type. As a comparison, the simulation for single swirl flow is also performed. In these simulations, the particles are introduced in the vicinity of pipe inner wall. In addition, the assumption that particles follow the flow motion absolutely is used. From the results, the motion of particles in these three cases is completely different. For the co-rotating and counter-rotating type, the particle entrainment into the main axial flow is clearly observed. This indicates the mixing is improved compared to single swirl flow. The difference of particle entrainment motion between co-rotating and counter-rotating type is slight.

DOI： 10.1007/s12206-015-0209-8

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Language：English   Publishing type：Research paper (international conference proceedings)

DOI： 10.1007/978-3-319-11541-2_24

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Language：Japanese   Publishing type：Research paper (scientific journal)

With the multi-combustion technology, the combustion in a hybrid rocket engine (HRE) can be temporarily stopped via oxidizer throttling. In this paper, two types of HREs, one with multi-combustion technology and the other without, are compared to investigate the effects of multi-combustion on the flight performance of launch vehicles (LVs). Non-dominated Sorting Genetic Algorithm-II (NSGA-II) which was a multi-objective evolutionary algorithm (MOEA) was applied to solve the design problems using real-number coding and the Pareto ranking method. To investigate the effects of the multi-combustion on flight performance of LV with HRE, three design problems were considered. The first case was the maximization of the flight altitude and the minimization of the gross weight. The second case was the minimization of the maximum acceleration and the minimization of the gross weight. The final case was the maximization of the flight downrange and the minimization of the gross weight. Many non-dominated solutions were obtained by NSGA-II, and a trade-off was observed between the two objective functions in each case. MOEA results were visualized using a parallel coordinate plot. According to the exploration result, it was found that the multi-combustion of HRE was effective to reduce the maximum acceleration. Such ability could be expected to reduce the shock load to payloads carried by the LV with HRE.

DOI： 10.11394/tjpnsec.6.137

CiNii Article

Other Link： https://ci.nii.ac.jp/naid/130005114801

Language：English   Publishing type：Research paper (international conference proceedings)

As a technique to eliminate O/F shifts, Altering-intensity Swirling Oxidizer Flow Type (A-SOFT) hybrid rocket is proposed using the combination axial and tangential oxidizer injection. In this paper, the possible geometric design point of the motor is considered about 2-tons class single stage A-SOFT hybrid rockets, and the nominal performance increase by eliminating O/F shifts is evaluated. The highest altitude of the A-SOFT is 450 [km] and the averaged ISP is 284[s] with 100[kg] payload, and this performance is several percent higher than the ones of S-520 with 95[kg] payload, Japanese 2-tons class solid sounding rocket. Though A-SOFTs increase the nominal flight performance by 1% from Swirling Oxidizer Flow Type hybrid rockets, which cannot O/F shift, their feedback control prevents the large performance losses caused by fuel regression errors and residuals, and this function is not found in the conventional hybrid rockets. In this paper, a planning of steady state burning tests of A-SOFT bread board model is also explained. The purpose of the burning tests are to demonstrate the concept of A-SOFTs. The test motor is 250[N] scale and uses polypropylene and GOX and its burning duration is 5[s]. In more than 10 times steady state tests, 50% throttling conditions with effective geometric swirl intensity between 0 and 37.3 are inculded.

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Language：English   Publishing type：Research paper (international conference proceedings)

DOI： 10.1016/j.proeng.2014.12.526

DOI： 10.1016/j.proeng.2014.12.526

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DOI： 10.1299/jfst.2014jfst0071

DOI： 10.1299/jfst.2014jfst0071

Other Link： https://www.jstage.jst.go.jp/article/jfst/9/5/9_2014jfst0071/_article/-char/en

Language：English   Publishing type：Article, review, commentary, editorial, etc. (scientific journal)

DOI： 10.1007/978-3-319-13359-1_39

Language：Japanese   Publishing type：Research paper (scientific journal)

DOI： 10.1299/transjsme.2014trans0287

Kyutacar

Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1299/jamdsm.2014jamdsm0023

DOI： 10.1299/jamdsm.2014jamdsm0023

Other Link： https://www.jstage.jst.go.jp/article/jamdsm/8/3/8_2014jamdsm0023/_article/-char/en

Language：Japanese   Publishing type：Research paper (scientific journal)

DOI： 10.2322/astj.13.41

DOI： 10.2322/astj.13.41

Authorship：Lead author,　Corresponding author   Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.2322/tastj.12.Pa_15

DOI： 10.2322/tastj.12.Pa_15

Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2014005382

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2014005382

Other Link： https://www.dl.begellhouse.com/journals/17bbb47e377ce023,76c2cb2847ba870e,031c13d34dedc6bd.html

Language：Japanese   Publishing type：Research paper (scientific journal)

DOI： 10.11394/tjpnsec.4.85

DOI： 10.11394/tjpnsec.4.85

Authorship：Lead author,　Corresponding author   Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2013005406

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2013005406

Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1299/spacee.6.15

DOI： 10.1299/spacee.6.15

Other Link： https://www.jstage.jst.go.jp/article/spacee/6/1/6_15/_article/-char/en

Language：English   Publishing type：Research paper (scientific journal)

Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1299/jtst.8.380

DOI： 10.1299/jtst.8.380

Other Link： https://www.jstage.jst.go.jp/article/jtst/8/2/8_380/_article/-char/en

Language：English   Publishing type：Research paper (scientific journal)

Language：English   Publishing type：Research paper (scientific journal)

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2012001351

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2012001351

Authorship：Lead author   Language：Japanese   Publishing type：Research paper (scientific journal)

液体酸素（LOX）を用いた酸化剤流旋回型ハイブリッドロケットエンジンでは，インジェクターより上流でLOXを気化する必要がある．筆者らは，LOXを気化する方法の一つとして，再生冷却方式のLOX気化ノズルを提案している．本研究では，推力1500N用のLOX気化ノズルを設計製作し，設計値より低い酸素流量と燃焼室圧条件において独立気化方式と再生冷却方式による気化燃焼実験を行った．独立気化方式の気化実験によって，LOXの気化とノズルの安全性が確認され，数値計算によるLOX気化ノズルの設計が適切であることが分かった．再生冷却方式の気化燃焼実験では，確実な着火と安定した燃焼が得られ，LOX気化ノズルを用いた酸化剤流旋回型ハイブリッドロケットエンジンの自立燃焼に成功した．また，LOXを気化させることで，LOXに直接旋回を与える場合より燃料後退速度やC*効率を向上できた．

DOI： 10.2322/stj.6.47

Kyutacar

Language：English   Publishing type：Research paper (international conference proceedings)

ドイツ   ハイデルベルグ  

An experimental study was performed on the combustion characteristics of a jet diffusion flame of Mg vapor injected through a small nozzle into CO2 atmospheres at low pressures from 8 to 48 kPa with a view to using Mg as fuel for a CO2-breathing turbojet engine in the Mars atmosphere. The Mg vapor jet produced three types of the flame. At lower pressures and higher injection velocities, a red-heated jet flame formed, in which the injected Mg vapor was heated by spontaneous reactions, turning red. At medium pressures and injection velocities, a stable luminous lifted-like flame developed above the rim of the chimney, a tube-like combustion product for the Mg vapor passage that grew on the nozzle during combustion. The flame had similar flame length properties to laminar jet diffusion flames of gaseous fuels. At higher pressures and lower injection velocities, a stable luminous attached flame developed at the rim of the chimney. The same reactions, producing MgO(g), CO and MgO(c), proceeded preferentially for all flames and chimneys. Carbon was only subordinately generated. Burning behavior of Mg vapor jets in a CO2 atmosphere has been represented, including the homogeneous reaction of Mg vapor with CO2, the diffusion of CO2, and the condensation and deposit of MgO. The injection velocity of Mg vapor at the rim of the chimney and the exothermic reactions with diffused CO2 that occur there play a crucial role in the attachment and development of the flames. The flame structure may be explained in terms of the relatively low gas-phase reaction rate of Mg with CO2.

DOI： 10.1016/j.proci.2006.07.192

Authorship：Lead author   Language：Japanese   Publishing type：Research paper (scientific journal)

We have proposed a swirling oxidizer type hybrid rocket engine. In this paper, liquid oxygen (LOX) was used as oxidizer. Combustion tests of a hybrid rocket engine with a swirling LOX flow were conducted by changing the swirl strength. Ignition was rapid and reliable, and combustion of PMMA with swirling LOX was stable. Fuel regression rates, C* efficiency and specific impulse of the hybrid rocket engine with swirling LOX flow were smaller than those with swirling gaseous oxygen (GOX). This low performance may be restraint of atomization and vaporization of LOX by formation of a liquid layer on the PMMA fuel and a decline of angular momentum of the swirling LOX during vaporization. Combustion oscillation occurred when the ratios of differential pressure between injector pressure and chamber pressure to chamber pressure were small. This combustion oscillation was confirmed to be a “Chugging” mode due to combustion time lag of LOX.

DOI： 10.2322/jjsass.54.242

Kyutacar

Language：Japanese   Publishing type：Article, review, commentary, editorial, etc. (scientific journal)

Kyutacar

Language：English   Publishing type：Research paper (scientific journal)

To develop an experimental small hybrid rocket with a swirling gaseous oxygen flow type engine, we made a flight model engine. Burning tests of the engine showed that a maximum thrust of 692 N and a specific impulse of 263 s (at sea level) were achieved. We designed a small hybrid rocket with this engine. The rocket measured 1.8 m in length and 15.4 kg in mass. To confirm the flight stability of the rocket, wind tunnel tests using a 112-scale model of the rocket and simulations of the flight attitude and trajectory were carried out. A flight test was conducted at Taiki-cho, Hokkaido, Japan on March 2001. The rocket reached an altitude of about 600 m, thus recording the first successful flight of a hybrid rocket in Japan. For the next stage, future issues to develop larger hybrid rockets using a swirling liquid oxygen flow type engine are discussed, and preliminary burning tests of the engine have been carried out.

DOI： 10.11230/jsts.21.1_1

Kyutacar

Event date： 2023.11.09 - 2023.11.10   Language：Japanese   Country：Japan  

Event date： 2023.11.06 - 2023.11.08   Language：English  

Event date： 2023.10.17 - 2023.10.20   Language：Japanese   Country：Japan  

Event date： 2023.07.03 - 2023.07.09   Language：English   Country：Japan  

Event date： 2023.07.03 - 2023.07.09   Language：English   Country：Japan  

Event date： 2023.07.03 - 2023.07.09   Language：English   Country：Japan  

Event date： 2023.02.14   Language：Japanese   Country：Japan  

Event date： 2023.02.14   Language：Japanese   Country：Japan  

Event date： 2022.11.10 - 2022.11.11   Language：Japanese  

Event date： 2022.11.09 - 2022.11.11   Language：English  

Event date： 2022.11.01   Language：Japanese   Country：Japan  

Event date： 2022.02.28 - 2022.03.04   Language：English   Country：Japan  

Event date： 2022.02.28 - 2022.03.04   Language：English   Country：Japan  

Event date： 2022.02.14   Language：Japanese  

Event date： 2022.02.14   Language：Japanese  

Event date： 2022.01.13 - 2022.01.14   Language：Japanese  

Event date： 2021.11.19   Language：Japanese  

Event date： 2020.11.27   Language：Japanese  

Event date： 2020.10.28 - 2020.10.30   Language：English