担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）

<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

その他リンク： https://ci.nii.ac.jp/naid/130008034521

記述言語：英語   掲載種別：研究論文（学術雑誌）

Burning tests of a laboratory-scale hybrid rocket engine were carried out with gaseous oxygen and a microcrystalline-wax-based fuel to look into the feasibility of using an intrusive resistor-based sensor for measuring the fuel regression rate. This initial screening was driven by the need for real-time control of the oxidizer-to-fuel ratio in altering-intensity swirling-flow-type hybrid rocket engines aiming at performance optimization. A traditional ballistic reconstruction technique was critically revised in order to build up a framework for comparison with the measured data; with the measured aft-chamber pressure and oxygen mass flow rate time histories, the fuel regression rate and port diameter were reconstructed over the firing by estimating the combustion efficiency with the constraint that calculated and measured fuel mass consumed are equal. This technique invariably suffers from the issue of presenting multiple solutions for the fuel mass flow rate in the proximity of the optimum mixture ratio, for which a novel variable-efficiency approach is proposed. Reconstructed data show that regression rate is nearly constant in each firing, yielding dependence upon the port diameter other than the mass flux. Resistor-sensor raw data displayed large deviation from the ballistic results for the slower burning rate of the sensor support. A detailed analysis is presented.

DOI： 10.2514/1.A34695

DOI： 10.2514/1.A34695

Scopus

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記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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

Scopus

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記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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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記述言語：英語   掲載種別：研究論文（学術雑誌）

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

Scopus

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担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）

<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

その他リンク： https://ci.nii.ac.jp/naid/130007642475

記述言語：英語   掲載種別：研究論文（学術雑誌）

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.

DOI： 10.1299/jfst.2018jfst0029

Scopus

CiNii Article

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担当区分：筆頭著者   記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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.

Scopus

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記述言語：英語   掲載種別：研究論文（学術雑誌）

A multidisciplinary design methodology for a launch vehicle (LV) with a hybrid rocket was proposed in order to obtain qualitative design knowledge. In this study, a propulsion system performance and aerodynamic performance were empirically evaluated using a theoretical approach, and flight evaluation was performed by solving equations of motion to evaluate the motion along the horizontal/vertical directions and rotation around the body axis. To demonstrate the applicability of the proposed method, two types of multiobjective design problems were solved using multiobjective evolutionary algorithms, and the results were visualized using the parallel coordinate plot, which is a data mining technique. The first design problem has two objective functions: the maximization of the highest altitude and the minimization of the total mass. In this case, the effect of the constraint related to the body length and diameter ratio was evaluated. The second design problem also had two objective functions and two constraints for a realistic LV design: the maximization of the downrange and the minimization of the total mass. The target altitude was different for these problems although the objective functions were identical. As a result, the trade-off information was successfully acquired for each design problem. It was also found that the body length and diameter ratio were key factors in deciding the maximum altitude. Furthermore, during the downrange maximization, the characteristic geometries observed in the obtained nondominated solutions revealed the design required to fulfill the constraint related to the target altitude.

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

Scopus

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記述言語：英語   掲載種別：研究論文（学術雑誌）

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

Scopus

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担当区分：筆頭著者   記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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

Scopus

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記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1504/IJAL.2016.074912

記述言語：英語   掲載種別：研究論文（学術雑誌）

This paper reports on the conceptual design of a three-stage launch vehicle (LV) with a clustered hybrid rocket engine (HRE) through multi-disciplinary design optimization. This LV is a space transportation concept that can deliver micro-satellites to sun-synchronous orbits (SSOs). To design a high-performance LV with HRE, the optimum size of each component, such as an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank, and a nozzle, should be determined. In this study, paraffin (FT-0070) is used as a propellant for the HRE, and three cases are compared: In the first case, HREs are optimized for each stage. In the second case, HREs are optimized together for the first and second stages but separately for the third stage. In the third case, HREs are optimized together for each stage. The optimization results show that the performance of the design case that uses the same HREs in all stages is 40% reduced compared with the design case that uses optimized HREs for each stage.

DOI： 10.1299/jfst.2016jfst0003

CiNii Article

その他リンク： https://ci.nii.ac.jp/naid/130005126936

担当区分：筆頭著者   記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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

Scopus

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担当区分：筆頭著者   記述言語：英語   掲載種別：研究論文（学術雑誌）

At the Hybrid Rocket research Working Group, performing demonstrations of current technologies is pursued. The target thrust was set as 5 kN and a hybrid rocket test engine program (HTE-5-1) was initiated. To demonstrate the hybrid rocket engine technology, a hybrid rocket test facility was designed at the Akiruno facility, which belongs to the Japan Aerospace Exploration Agency. A low-flow rate gaseous oxygen (GOX) feed system that includes a purge system, test stand, measurement system, and control system was manufactured. The operability and functionality, such as the GOX mass flow range and natural frequency of the test stand, were examined. It was confirmed that combustion tests under a condition of up to 0.5 kg/s GOX mass flow rate could be conducted. A 5-kN-thrust swirling-oxidizer-flow-type hybrid rocket engine using a GOX/polypropylene propellant was designed and manufactured as a demonstrator. Two combustion tests were performed with the demonstration engine under conditions below the design point. Reliable ignition and stable combustion for long durations were achieved, and accurate experimental data were obtained. This demonstrated the effectiveness of the combustion test and validated the suitability of our engine design method. It was confirmed that the combustion test could be performed safely and properly using the new test facility and demonstrator.

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2017011994

Scopus

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記述言語：英語   掲載種別：研究論文（学術雑誌）

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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記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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.

Scopus

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記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

A single-stage launch vehicle with hybrid rocket engine has been conceptually designed by using design informatics, which has three points of view, i.e., problem definition, optimization, and data mining. The primary objective of the present design is that the downrange and the duration time in the lower thermosphere are sufficiently secured for the aurora scientific observation, whereas the initial gross weight is held down to the extent possible. The multidisciplinary design optimization was performed by using a hybrid evolutionary computation. Data mining was also implemented by using a scatter plot matrix. Polypropylene and liquid oxygen with swirling flow are adopted as solid fuel and liquid oxidizer, respectively. The condition of two-time ignitions is assumed in fight sequence on the equation of motion for the three degree of freedom rigid body. Consequently, the design information regarding the tradeoffs, the behaviors of the design variables in the design space to become the nondominated solutions, and the implication of the design variables for the objective functions have been obtained quantitatively. The structurization and visualization of the design space has been implemented in order to observe the effectiveness of the local regions of each design variable. The advantage of extinction-reignition has been indicated.

DOI： 10.2514/6.2016-0583

Scopus

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記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.20910/IJASE.2015.2.2.029

記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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

Scopus

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記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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

Scopus

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担当区分：責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）

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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記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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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記述言語：日本語   掲載種別：研究論文（学術雑誌）

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

その他リンク： https://ci.nii.ac.jp/naid/130005114801

記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

In this study, a multi-objective genetic algorithm (MOGA) was applied to the multidisciplinary design optimization (MDO) of a hybrid rocket. A swirling-oxidizer-type hybrid rocket engine (HRE) with a single cylindrical grain port was designed. It was considered that this HRE could temporarily stop combustion via oxidizer throttling; this feature is called multi-combustion. The MOGA was applied to solve the multi-objective problem using real-number coding and the Pareto ranking method. In this study, three design problems were considered. First problem was the maximization of the flight altitude and minimization of the gross weight. Second problem was the minimization of the maximum acceleration and minimization of the gross weight. Third problem was the maximization of the duration time over the target flight altitude and minimization of the gross weight. Each objective function was empirically estimated. In addition, this study compared two types of HREs to investigate the emects of the multi-combustion: one type was able to carry out the multi-combustion, and the other was not. Many non-dominated solutions were obtained using the MOGA, and a trade-off was observed between the two objective functions. To understand the design problem, the MOGA results were visualized using a parallel coordinate plot (PCP).

DOI： 10.1016/j.proeng.2014.12.526

DOI： 10.1016/j.proeng.2014.12.526

Scopus

その他リンク： https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=84978173389&origin=inward

記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

A single-stage launch vehicle with hybrid rocket engine has been conceptually designed by using design informatics, which has three points of view as problem definition, optimization, and data mining. The primary objective of the design in the present study is that the sufficient down range and the duration time in the lower thermosphere are achieved for aurora scientific observation whereas the initial gross weight is held down. Multidisciplinary design optimization and data mining were performed by using evolutionary hybrid computation under the conditions that polypropylene as solid fuel and liquid oxygen as liquid oxidizer were adopted and that single-time ignition is implemented in sequence. Consequently, the design information regarding the tradeoffs and the behaviors of the design variables in the design space was obtained in order to quantitatively differentiate the advantage of hybrid rocket engine.

DOI： 10.1007/978-3-319-11541-2_24

Scopus

その他リンク： https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=84962821808&origin=inward

記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1299/jfst.2014jfst0071

DOI： 10.1299/jfst.2014jfst0071

その他リンク： https://www.jstage.jst.go.jp/article/jfst/9/5/9_2014jfst0071/_article/-char/en

記述言語：英語   掲載種別：記事・総説・解説・論説等（学術雑誌）

DOI： 10.1007/978-3-319-13359-1_39

記述言語：日本語   掲載種別：研究論文（学術雑誌）

DOI： 10.1299/transjsme.2014trans0287

Kyutacar

記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1299/jamdsm.2014jamdsm0023

DOI： 10.1299/jamdsm.2014jamdsm0023

その他リンク： https://www.jstage.jst.go.jp/article/jamdsm/8/3/8_2014jamdsm0023/_article/-char/en

記述言語：日本語   掲載種別：研究論文（学術雑誌）

DOI： 10.2322/astj.13.41

DOI： 10.2322/astj.13.41

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.2322/tastj.12.Pa_15

DOI： 10.2322/tastj.12.Pa_15

記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2014005382

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2014005382

その他リンク： https://www.dl.begellhouse.com/journals/17bbb47e377ce023,76c2cb2847ba870e,031c13d34dedc6bd.html

記述言語：日本語   掲載種別：研究論文（学術雑誌）

DOI： 10.11394/tjpnsec.4.85

DOI： 10.11394/tjpnsec.4.85

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2013005406

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2013005406

記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1299/spacee.6.15

DOI： 10.1299/spacee.6.15

その他リンク： https://www.jstage.jst.go.jp/article/spacee/6/1/6_15/_article/-char/en

記述言語：英語   掲載種別：研究論文（学術雑誌）

記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1299/jtst.8.380

DOI： 10.1299/jtst.8.380

その他リンク： https://www.jstage.jst.go.jp/article/jtst/8/2/8_380/_article/-char/en

記述言語：英語   掲載種別：研究論文（学術雑誌）

記述言語：英語   掲載種別：研究論文（学術雑誌）

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2012001351

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2012001351

記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）

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

担当区分：筆頭著者   記述言語：日本語   掲載種別：研究論文（学術雑誌）

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

担当区分：筆頭著者   記述言語：日本語   掲載種別：研究論文（学術雑誌）

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

記述言語：日本語   掲載種別：記事・総説・解説・論説等（学術雑誌）

Kyutacar

記述言語：英語   掲載種別：研究論文（学術雑誌）

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

開催期間： 2023年11月09日 - 2023年11月10日   記述言語：日本語   開催地：福岡市   国名：日本国  

開催期間： 2023年11月06日 - 2023年11月08日   記述言語：英語   開催地：Sendai, Miyagi  

開催期間： 2023年10月17日 - 2023年10月20日   記述言語：日本語   開催地：熊本市   国名：日本国  

開催期間： 2023年07月03日 - 2023年07月09日   記述言語：英語   開催地：久留米市   国名：日本国  

開催期間： 2023年07月03日 - 2023年07月09日   記述言語：英語   開催地：久留米市   国名：日本国  

開催期間： 2023年07月03日 - 2023年07月09日   記述言語：英語   開催地：久留米市   国名：日本国  

開催期間： 2023年02月14日   記述言語：日本語   開催地：宇宙科学研究所、神奈川県相模原市 ZOOM   国名：日本国  

開催期間： 2023年02月14日   記述言語：日本語   開催地：宇宙科学研究所、神奈川県相模原市 ZOOM   国名：日本国  

開催期間： 2022年11月10日 - 2022年11月11日   記述言語：日本語   開催地：オンライン（Webex）  

開催期間： 2022年11月09日 - 2022年11月11日   記述言語：英語   開催地：Sendai, Miyagi  

開催期間： 2022年11月01日   記述言語：日本語   開催地：熊本市   国名：日本国  

開催期間： 2022年02月28日 - 2022年03月04日   記述言語：英語   開催地：オンライン（大分）   国名：日本国  

開催期間： 2022年02月28日 - 2022年03月04日   記述言語：英語   開催地：オンライン（大分）   国名：日本国  

開催期間： 2022年02月14日   記述言語：日本語   開催地：宇宙科学研究所、神奈川県相模原市 ZOOM  

開催期間： 2022年02月14日   記述言語：日本語   開催地：宇宙科学研究所、神奈川県相模原市 ZOOM  

開催期間： 2022年01月13日 - 2022年01月14日   記述言語：日本語   開催地：宇宙科学研究所、神奈川県相模原市 ZOOM  

開催期間： 2021年11月19日   記述言語：日本語   開催地：オンライン（ZOOM）  

開催期間： 2020年11月27日   記述言語：日本語   開催地：宇宙科学研究所、神奈川県相模原市  

開催期間： 2020年10月28日 - 2020年10月30日   記述言語：英語   開催地：Sendai, Miyagi