Faculty Profiles - KITAGAWA Koki

Construction of the Calculation Model to Clarify the Energy Conversion Ratio in the Entire Area of an LSD Wave Reviewed

ARAI Ten, MATSUI Kohei, KITAGAWA Koki

Journal of Evolving Space Activities ( International Symposium on Space Technology and Science ) 2 ( 0 ) n/a 2024.01

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

In laser propulsion, there is a method that uses the atmosphere as a propellant and laser energy from the ground. This method utilizes Laser Supported Detonation (LSD) in which a laser is absorbed and thrust is generated by detonation. To estimate the thrust of laser propulsion with LSD, the energy conversion ratio is an important parameter. The energy conversion ratio is the ratio of energy conversion from laser to working fluid. In a previous study, the energy conversion ratio was obtained experimentally and analytically, but the results have a large discrepancy. To eliminate the discrepancy, it is necessary a new calculation method for the energy conversion ratio. The purpose of this study is to establish a new method for calculating the energy conversion ratio across the entire LSD wavefront using the specific heat ratio of LSD obtained theoretically under the assumption of the Chapman-Jouguet state, and to evaluate the calculation method. The results show that the energy conversion ratios are 0.25 and 0.38 at the termination time under the laser energy of 8 and 10 J, respectively, which are close to the experimental values. This agreement could indicate that the newly proposed method for calculating energy conversion ratios is useful in laser propulsion.

DOI： 10.57350/jesa.191

Evaluation of Radiative Heat-Transfer Effect on Paraffin-Wax Regression Rate in Hybrid Rocket Reviewed International journal

Naka G., Carmicino C., Messineo J., Kitagawa K., Shimada T.

Journal of Propulsion and Power 39 ( 2 ) 258 - 273 2023.03

Function Test of Experimental System to Obtain Laser Ignition Characteristics of Low-Temperature Boron/Potassium Nitrate Reviewed International journal

MATSUI Kohei, MATSUURA Yoshiki, TOKUDOME Shinichiro, KITAGAWA Koki

Journal of Evolving Space Activities ( International Symposium on Space Technology and Science ) 1 ( 0 ) n/a 2023.01

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

In order to use laser ignition systems for solid rocket motors operating in deep space environments, it is necessary to elucidate the laser ignition characteristics of the ignition charge in low-temperature environments. This study aims to design an experimental system that can confirm the ignition threshold, ignition delay, and ignition temperature by irradiating an ignition charge with a diode laser in a low-temperature environment. Ignition experiments at room temperature were conducted. The data were evaluated statistically to obtain an ignition threshold with the maximum likelihood method. The relationship between the laser irradiation duration and the laser power with respect to the ignition threshold was obtained. The target value of the low-temperature environment temperature was determined as -50 °C. We examined the requirements of the experimental system and conceptually designed the system to simulate the low-temperature environment. It was confirmed that the constructed experimental system cooled the ignition charge to -50 °C. In the ignition experiment, the ignition charge was successfully ignited at room temperature and at low temperature. The ignition delay, the ignition temperature, and the high-speed image were obtained. Eventually, the validity of the experimental system was confirmed through the function tests.

DOI： 10.57350/jesa.17

Evaluation of Measurement System for Heat Transfer Characteristics ofLiquid Oxygen Flow in a Thin Tube under High Heat Flux Conditions Reviewed International journal

KITAGAWA Koki, MATSUI Kohei

Journal of Evolving Space Activities ( International Symposium on Space Technology and Science ) 1 ( 0 ) n/a 2023.01

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

Heat transfer characteristics of liquid oxygen under high pressure and high heat flux should be elucidated to design a regenerative cooling nozzle for a hybrid rocket. Previously, we developed an experimental system using Joule heating to a thin tube in which liquid oxygen flows. The objective of this study is the evaluation of the measurement system to obtain the heat transfer characteristic under high pressure and high heat flux. We conducted a cold flow experiment and an experiment of heat transfer characteristic measurement. The friction factor of the tube is estimated with the Colebrook-White equation. The heat transfer coefficient was calculated by solving the heat conduction equation in the tube and the conservation laws of the fluid. The calculated heat transfer coefficient is in the same order as the experimental value. Consequently, we confirmed that experimental and analytical methods are appropriate to obtain heat transfer characteristics.

DOI： 10.57350/jesa.45

Performance Evaluation of the Launch Vehicle Using WAX-Based Altering-Intensity Swirling-Oxidizer-Flow-Type Hybrid Rocket Engine by Means of Evolutionary Algorithms

KANAZAKI Masahiro, IMAMURA Taro

The Proceedings of OPTIS ( The Japan Society of Mechanical Engineers ) 2022.14 ( 0 ) U00073 2022.01

Language：Japanese Publishing type：Research paper (conference, symposium, etc.)

To improve the thrust by the O/F shift, which is one of the weak points of the hybrid rocket, a method with a variable oxidizer strength has been proposed. In this study, the superiority of this type of hybrid rocket over the conventional hybrid rocket without variable oxidizer supply was compared through several launch vehicle designs, such as altitude maximization by vertical launch and downrange extension problems with an optimized launch angle. A multi-objective evolutionary algorithm and the multi-disciplinary evaluation method of the launch vehicle were applied.

DOI： 10.1299/jsmeoptis.2022.14.u00073

Flight Results of Solid Propulsion Systems for Epsilon-4 Reviewed International journal

KITAGAWA Koki, TOKUDOME Shinichiro, HORI Keiichi, UI Kyoichi, KINOSHITA Masahiro, HASHIMOTO Junichi, ICHIMURA Kotaro

TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN ( THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES ) 19 ( 3 ) 400 - 406 2021.01

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

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.

DOI： 10.2322/tastj.19.400

Other Link： https://ci.nii.ac.jp/naid/130008034521

Development Results of Laser Ignition System for Solid Rocket Motor Reviewed International journal

MINAMI Keisuke, MATSUURA Yoshiki, KITAGAWA Koki, ARAKAWA Satoshi, MORISHITA Naoki, TAKEMAE Toshiaki, IWABUCHI Shota, WADA Asato, TOKUDOME Shinichiro

TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN ( THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES ) 19 ( 5 ) 807 - 811 2021.01

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

This paper describes the development status of a laser ignition system for a solid rocket motor. This system is being developed as a simple, lightweight, and small design with a high resistance to electrical disturbances and a high level of safety. The most notable advantage of this system is that its high level of safety can decrease the cost of launching rockets into space. A laser initiator and a laser safe-and-arm device (laser S/A), which are essential components of the proposed system, were developed. In particular, prototypes of the laser initiator and laser S/A for the ignition of an upper stage rocket motor were manufactured, and some environmental tests, which are required for space rocket devices, were conducted. In addition, the lowest laser energy that is needed to ignite the laser initiator was determined by changing the laser power and operating time of the laser S/A. Furthermore, a small rocket motor vacuum fire test was successfully conducted.

DOI： 10.2322/tastj.19.807

Other Link： https://ci.nii.ac.jp/naid/130008084441

Reconstructed Ballistic Data Versus Wax Regression-Rate Intrusive Measurement in a Hybrid Rocket Reviewed International journal

Jerome Messineo, Koki Kitagawa, Carmine Carmicino, Toru Shimada, Christian Paravan

Journal of Spacecraft and Rockets ( American Institute of Aeronautics and Astronautics ) 57 ( 6 ) 1295 - 1308 2020.07

Prediction of space and time distribution of wax-based fuel regression rate in a hybrid rocket Reviewed

Naka G., Messineo J., Kitagawa K., Carmicino C., Shimada T.

AIAA Propulsion and Energy 2020 Forum 1 - 30 2020.01

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

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

Other Link： https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85091289443&origin=inward

Development and Flight Results of Solid Propulsion System for Enhanced Epsilon Launch Vehicle Reviewed International journal

KITAGAWA Koki, TOKUDOME Shinichiro, HORI Keiichi, TANNO Haruhito, NAKANO Nobuyuki

TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN ( THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES ) 17 ( 3 ) 289 - 294 2019.01

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

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

DOI： 10.2322/tastj.17.289

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

Hybrid rocket firing experiments at various axial–tangential oxidizer-flow-rate ratios Reviewed

Ozawa K., Kitagawa K., Aso S., Shimada T.

Journal of Propulsion and Power 35 ( 1 ) 1 - 15 2019.01

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

Other Link： https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85076482957&origin=inward

Evaluation of safety distance for blast of hybrid rocket propellants Reviewed

Takahashi A., Kitagawa K., Shimada T.

AIAA Propulsion and Energy Forum and Exposition, 2019 2019.01

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

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

Other Link： https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85095967657&origin=inward

Image analysis for velocity profile estimation in A-SOFT hybrid rocket combustor Reviewed International journal

Kimura N., Obata K., Kitagawa K., Shimada T.

Journal of Fluid Science and Technology ( The Japan Society of Mechanical Engineers ) 13 ( 4 ) JFST0029 - JFST0029 2018.01

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

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

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Flight results of solid propulsion system for epsilon launch vehicle from the third flight Reviewed

Kitagawa K., Tokudome S., Hori K., Ui K., Kinoshita M., Hashimoto J., Ichimura K.

Proceedings of the International Astronautical Congress, IAC 2018-October 2018.01

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.

Other Link： https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85065301986&origin=inward

Design Methodology of a Hybrid Rocket-Powered Launch Vehicle for Suborbital Flight Reviewed

Kanazaki M., Yoda H., Chiba K., Kitagawa K., Shimada T.

Journal of Aerospace Engineering 30 ( 6 ) 2017.11

Performance and regression rate characteristics of 5-kN swirling-oxidizer-flow-Type hybrid rocket engine Reviewed

Sakurai T., Yuasa S., Ando H., Kitagawa K., Shimada T.

Journal of Propulsion and Power 33 ( 4 ) 891 - 901 2017.01

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

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

Other Link： https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85021289537&origin=inward

Development of solid propulsion system for enhanced epsilon launch vehicle and epsilon's second launch results Reviewed

Kitagawa K., Tokudome S., Hori K., Hashimoto J., Nakano N., Tanno H.

Proceedings of the International Astronautical Congress, IAC 13 8590 - 8596 2017.01

Structurisation and visualisation of design space for launch vehicle with hybrid rocket engine Reviewed International journal

Kazuhisa Chiba; Masahiro Kanazaki; Shin'ya Watanabe; Koki Kitagawa; Toru Shimada

International Journal of Automation and Logistics 2 ( 1/2 ) 26 - 44 2016.02

Design optimization of launch vehicle concept using cluster hybrid rocket engine for future space transportation Reviewed

KANAZAKI Masahiro, ITO Shoma, KANAMORI Fumio, NAKAMIYA Masaki, KITAGAWA Koki, SHIMADA Toru

Journal of Fluid Science and Technology 11 ( 1 ) JFST0003 - JFST0003 2016.01

Static burning tests on a bread board model of altering-intenisty swirling-oxidizer-flow-type hybrid rocket engine Reviewed

Ozawa K., Usuki T., Mishima G., Kitagawa K., Yamashita M., Mizuchi M., Katakami K., Maji Y., Aso S., Tani Y., Wada Y., Shimada T.

52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 2016.01

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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Effect of oxidizer particle orientation on burning rates of composite propellants Reviewed

Hasegawa H., Fukunaga M., Kitagawa K., Shimada T.

International Journal of Energetic Materials and Chemical Propulsion 15 ( 4 ) 285 - 304 2016.01

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

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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Development of test facilities for 5 kn-thrust hybrid rocket engines and a swirling-oxidizer-flow-type hybrid rocket engine for technology demonstration Reviewed

Kitagawa K., Yuasa S., Sakurai T., Hatagaki S., Shiraishi N., Ando H., Yagishita T., Suzuki N., Takayama A., Yui R., Shimada T.

International Journal of Energetic Materials and Chemical Propulsion 15 ( 6 ) 435 - 451 2016.01

Development of solid propulsion system for enhanced epsilon launch vehicle and M-35 static firing test Reviewed

Kitagawa K., Tokudome S., Hori K., Tanno H., Nakano N.

Proceedings of the International Astronautical Congress, IAC 2016.01

Ascendancy of extinction-reignition on single-stage hybrid sounding rocket in view of fuels Reviewed

Chiba K., Yoda H., Ito S., Kanazaki M., Watanabe S., Kitagawa K., Shimada T.

57th AIAA/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference 2016.01

Design Optimization of Single-Stage Launch Vehicle Using Hybrid Rocket Engine,” International Journal of Aerospace System Engineering Reviewed International journal

Kanazaki, M., Ariyarit, A., Yoda, H., Ito, S., Chiba, K., Kitagawa, K., Shimada, T.

International Journal of Aerospace System Engineering 2 ( 2 ) 29 - 33 2015.12

Evolutionary algorithm applied to ballistic launch vehicle design using hybrid rocket engine evaluated by enhanced flight simulation Reviewed

Yoda H., Ito S., Kanazaki M., Chiba K., Kitagawa K., Shimada T.

2015 IEEE Congress on Evolutionary Computation, CEC 2015 - Proceedings 618 - 625 2015.09

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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Multidisciplinary design exploration for sounding launch vehicle using hybrid rocket engine in view of ballistic performance Reviewed

Chiba K., Kanazaki M., Ariyarit A., Yoda H., Ito S., Kitagawa K., Shimada T.

International Journal of Turbo and Jet Engines 32 ( 3 ) 299 - 304 2015.09

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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Numerical analysis of multi-parallelized swirling flow inside a circular pipe Reviewed

Takayama A., Kitagawa K., Shimada T.

Journal of Mechanical Science and Technology 29 ( 3 ) 951 - 962 2015.03

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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Conceptual design of single-stage launch vehicle with hybrid rocket engine using design informatics Reviewed

Chiba K., Kanazaki M., Kitagawa K., Shimada T.

Computational Methods in Applied Sciences 36 369 - 384 2015.01

Evaluation of Multi-Time-Combustion-Type Hybrid Rocket by Means of Multi-objective Evolutionary Algorithm Reviewed

Kanazaki Masahiro, Chiba Kazuhisa, Kitagawa Koki, Shimada Toru

Transaction of the Japanese Society for Evolutionary Computation ( The Japanese Society for Evolutionary Computation ) 6 ( 3 ) 137 - 145 2015.01

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

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

Performance calculations and burning tests on altering-intensity swirling oxidizer flow type hybrid rocket engines Reviewed

Ozawa K., Kitagawa K., Shimada T.

Proceedings of the International Astronautical Congress, IAC 9 7332 - 7345 2015.01

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.