野中 聡

© 2019 IEEE. We have been developing a liquid hydrogen (LH 2 ) cooled superconducting energy apparatus, such as superconducting generator, SMES, and so on. An MgB 2 superconductor whose critical temperature is 39 K is now developing for a practical use. It can be cooled by LH 2 with a sufficient temperature margin. An MgB 2 wire is expected to be used for superconducting equipment due to the low production cost and material cost. In order to design equipment using the MgB 2 wire, we have carried out measurement tests of critical superconducting properties of MgB 2 under LH 2 cooling and investigation of heat transfer characteristics of LH 2 . There are few reports on the experimental results of a superconducting coil using a long MgB 2 wire under the LH 2 immersion cooling. In this study, we have carried out an excitation test of a small MgB 2 coil immersed in LH 2 under an external magnetic field. Tanaka et al. proposed that the test coil is a 529 turn solenoid coil with an inner diameter of 120 mm, an outer diameter of 190 mm, and a height of 41 mm, which was produced by the Wind and React method using a 300-m multifilament MgB 2 wire (Hitachi Ltd., Ibaraki, Japan). The temperature of LH 2 was changed from 21 K to 30 K and the external magnetic field was also applied up to 4.5 T. In the experiment, the critical characteristic of the solenoid coil was measured and a coil load line was obtained.

© 2002-2011 IEEE. Developing applications of liquid-hydrogen (LH 2 )-cooled superconducting devices is a challenging issue. Since the boiling point of LH 2 is 20.4 K, MgB 2 , whose critical temperature is 39 K, can be cooled with a sufficient temperature margin. Furthermore, MgB 2 wire is expected to be used for superconducting equipment due to the low production and material cost. Therefore, in order to design MgB 2 superconducting equipment, the knowledge of normal zone propagation phenomena is important for the thermal stability and the quench protection. In this study, normal zone propagation and minimum quench energy (MQE) with a multi-filamentary MgB 2 superconducting wire produced by Hitachi, Ltd. were observed under immersed in LH 2 . In the experiment, heat pulse, which initiates a normal zone, was injected to the center area of 200 cm long MgB 2 wire. Then, the MQE and the normal zone propagation velocity (NZPV) were measured under specific conditions. NZPV was in the order of several cm/s and MQE was in the order of a few J at 30 K under LH 2 cooling. In order to clarify temperature distribution along the wire during the normal zone propagating phenomena, the simulation model of MgB 2 wire cooled by LH 2 was created and analyzed using a finite element method simulation software.

© 2019 Published under licence by IOP Publishing Ltd. The knowledge of heat transfer properties of liquid hydrogen is important for designing and developing superconducting devices. In this study, film boiling heat transfers of liquid hydrogen flowing inside of heated pipe were measured under saturated conditions at the absolute pressures of 700 kPa for various mass flow rates. Test pipe heater made of SS310S with a diameter of 8 mm and a length of 200 mm was used. The test heater was once heated up to the film boiling regime with exponential heat generation rate. And then, while the heat generation rate was decreased exponentially down to the minimum heat flux, the film boiling heat transfer coefficient, mass flow rate per unit area and the degree of superheat of pipe length direction were measured. It was observed that though the mass flow rate decreased according to increase of the heat generation rate, the heat transfer coefficient increased. Discussions on the experimental results of various conditions were carried out to clarify the phenomenon of film boiling of liquid hydrogen flowing inside of heated pipe. It was considered that since the void fraction in the flow path was high, the effect of improving the heat transfer rate was also observed in the film boiling region due to the acceleration of the liquid phase. © Published under licence by IOP Publishing Ltd.

© 2019 Published under licence by IOP Publishing Ltd. Film boiling heat transfer coefficients were measured for the test wires of 0.5 and 0.7 mm in diameters and 200mm in length located at the centre of 8 mm diameter conduit with upward flow of liquid hydrogen. Pressures are ranged from 0.4 to 1.1 MPa, liquid subcoolings from 0 to 11 K and flow velocities up to 5 m/s. The experimental data were compared with the authors' correlation for forced flow film boiling already presented based on the data for 1.2 mm diameter test wire. The experimental data are higher than the predicted values by the correlation. Discussion is made on the effect of wire diameter observed and a new correlation was presented. © Published under licence by IOP Publishing Ltd.

Copyright © 2019 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Space transportation vehicles have asymmetric protuberant devices on their surfaces, such as cable ducts. Such protuberances, typically arranged asymmetrically with respect to the vehicle axis, are known to cause asymmetric vortices and side force. In this study, to understand effects of side force and its associated flow fields, both wind-tunnel tests and numerical calculations for a slender body with an asymmetric protuberance were conducted at Mach 1.5. The results of computed aerodynamic coefficients are in good agreement with the experimental results and detailed flow structures are provided. In particular, the results revealed that side force was generated by two factors. It linearly increased as vortices detached from the body, and nonlinearly increased based on the effects of secondary vortex scale as the angle of attack increased. Additionally, the axial position and azimuthal angle (angle along the circumferential direction around the body axis) of the protuberance strongly influenced side force characteristics. First, the side force was significantly higher when the protuberance was installed in a forward axial position. Second, when the protuberance was installed on the leeward side of the slender body, the side force increased with the angle of attack. These results are not limited to the presented configuration but to other rocket designs.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. Technical demonstrations for a reusable sounding rocket have been successfully conducted from 2010 to 2016 in ISAS/JAXA, and most of the technical issues identified for development of an operation system have been demonstrated. We are technically ready for the next phase for a reusable rocket system development from the results of technical maturation studies, and an activity of a system level verification study by a flight demonstrator RV-X (Reusable Vehicle Experiment) is proposed and underway now as collaborative study with private companies. Design and development of RV-X vehicle is currently in progress by maximum use of existing components and technical outcomes obtained from RLV related studies in JAXA. To obtain characteristics of the engine and propellant system of RV-X, system level ground firing tests have been conducted in 2018. Two flight campaigns are planned in this flight demonstration study. In this paper, the present status of preparations for RV-X flight demonstrations are presented.

© 2019 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Aerodynamic characteristics of space transportation vehicles have been investigated in order to guarantee their flight feasibilities. Over their surface, various protuberant devices are located such as cable ducts, and the presence of the protuberances can create flow asymmetry which especially contributes to the side force. In this study, to understand the aerodynamic effects of protuberance positions systematically, we conducted numerical calculations using super computer and supersonic wind tunnel tests at the Japan Aerospace Exploration Agency (JAXA) on slender-bodied launch vehicle with a protuberance at Mach 1.5. According to these results, with angles of attack, the presence of the protuberance contributes to the formation of a new vortex, which is composed of two co-rotating vortices; the one is the leeside vortex from the body, the other is the wake of the protuberance. This merged vortex generates the asymmetry of the surface pressure which enlarges the side force. It is revealed that when the protuberance is located in front (upstream), the side force by this vortex significantly increases with the angle of attack up to 15 °. Additionally, when the protuberance is installed on the leeward side, the side force is larger than on the windward side.

<p>It is known that aerodynamic characteristics of a slender body vary substantially at high angles-of-attack (AoAs), and then, will have strong impacts on its flight. For, example, the yaw force makes flight unstable. In this study, we investigated the relation between the yaw force and the configuration, and details of flowfield around the slender-bodied-vehicle numerically. The configuration consisting of “nose cone” and “square aftbody” parts was employed as the baseline, and then, compared with other three configurations having different fineness ratios. According to our computed results, in the case of 50 degrees of AoA, the longer the model became, the more asymmetry appeared: yaw force and asymmetry were found to be attributed not only to the length of the body, but also to the nose bluntness. On the contrary, in the case of 140 degrees, the shorter the model became, the more asymmetry appeared. Furthermore, the large nose bluntness increased C_Y. Interestingly, this trend is totally opposite to that observed at 50 degrees. It had been considered that the large nose bluntness and the small fineness ratio can reduce asymmetry and C_Y, however, this study showed that it is not true in the case over 90 degrees, due to complex wake flow structure discovered in the present numerical simulations.</p>

<p>The development of a fully reusable vertical-takeoff-and-vertical-landing (VTVL) rocket is indispensable for reducing space transportation costs. However, there are many technical issues associated with such vehicles, such as turnover maneuvers during return flight where the pitching moment plays a key role. It is known that aerodynamic characteristics can be controlled by installing aerodynamic devices, but the relationship between the aerodynamic characteristics and the flowfields has not been explored. To clarify this relationship using computational fluid dynamics (CFD), we investigated these flowfields and aerodynamic characteristics, in the case where we install such devices (fins) in the nose part of a reusable rocket. We found that vortices form downstream of the aerodynamic devices. For angles of attack between 0 and 90 degrees (in which the fins are located in the upstream portion), these vortices significantly affect the surface pressure on the rocket and increase the pitching moment. On the other hand, for AOAs between 90 to 180 degrees (in which the fins are in the downstream portion), the effect of these vortices on the on-surface pressure is negligible, and only vortices formed near the surface of the fins increase the pitching moment.</p>

<p>Most of flight vehicles have various protuberant devices on their surfaces, but asymmetry in their positioning with respect to the body axis can affect aerodynamic characteristics of vehicles, particularly roll moment. Thus, it is important in rocket development to clarify the effects of the protuberances on the vehicle aerodynamic characteristics. In this study, as a basic research, we systematically investigated such effects using CFD, by changing the positions of a protuberance. As a result, the roll moment increased nearly linearly with angle of attack (=α), but its trend was different in protuberance locations, particularly when arranged near the center-of-gravity. In positioning there at α = 20 °, the wake vortex center moved farther away from protuberance compared with α = 15 °, then the pressure decline at its wake side was suppressed, and thus, the pressure difference between its upstream and downstream sides became smaller. As a consequence, the roll moment did not arise linearly, but decreased at α = 20 °.</p>

Copyright © 2018 by the authors. Torealize high-performance cryogenic propulsion systems, the chilldown sequence has to be improved. Because the chilldown is carried out under low gravity, the effect of gravity on the two-phase flow, especially at low flow rate, should be investigated. To understand the physics under low gravity, an experiment was conducted using a sounding rocket. Two identical test sections with different mass flow rates simulated part of a turbopump, each of which has a complex flowpath including slits and a dead end. Using liquid nitrogen, the flight experiment obtained data of temperatures, pressures, void fractions, and video frames of liquid motion. Then, the flight experiment data were compared to the ground data taken under normal gravity, revealing that the slits played an important role in the chilldown process and that the test sections were quickly chilled down under low gravity. The slits of the test sections formed liquid jets, and their behaviors were different from those in the ground experiment. In the flight experiment, the jets easily reached the dead end of the test sections and cooled down the whole walls due to the increase in inertia and wettability; however, such behaviors were hardly observed in the ground experiment. The difference between the ground and flight is significant at lower flow rate.

© 2019 SPIE. By using two high-speed cameras and a slightly extended visualization setup (typically based on a Toepler system) one can generate two simultaneous time-resolved records of the same flow, where these records can be obtained with different visualization methods, different spatial and different temporal resolutions. This allows one to generate visualizations that can complement each other in various ways and thus yield a considerably increased amount of information on the observed flow.

In order to make the access to space for scientific researches much easier and make the opportunities of the rocket launches much frequent, a fully reusable sounding rocket is proposed in ISAS/JAXA. The mission definition of the proposed reusable sounding rocket are 1) To achieve 100km in altitude and returns to the launch site, 2) The 100kg payload to be carried, 3) Flight frequency is higher than 10 times per a year, 4) The minimum flight interval is one day, and 5) Operational flight cost should be an order of magnitude less than the existing ISAS sounding rocket. Reusable sounding rocket is different from the present expendable rockets in 1) repeated operations, 2) returning flight / reignition of engine / vertical landing, that is, 3) fault tolerant / health management. Some key technologies related to these characteristics of reusable system have been verified to design an operative reusable sounding rocket in phase-A. Technologies verifications respect to the reusable vehicle, 1) reusable engine development and repeated operations, 2) reusable insulation development for cryogenic tank, 3) aerodynamic design and model flight demonstration for returning flight, 4) cryogenic liquid propellant management demonstration, 5) landing gear development, and 6) health management system construction, have been successfully conducted from 2010 to 2016.After these technical demonstrations, we are proceeding with a study for system level verifications by a flight demonstrator from 2016 as the next step for the development of reusable sounding rocket. In this plan, a small test vehicle will be established for repeated flight demonstrations. Objectives of the demonstration are 1) system architecture study for repeated flight operation such as quick turnaround operation and fault tolerant design method, 2) life controlled and frequently repeated use of cryogenic propulsion system and its flight demonstrations, 3) study for the advanced returning flight method of vertical landers and its flight demonstrations, and 4) demonstration of advanced technology for future RLVs such as more composite on board, in flight fuel management, GH2/GOX auxiliary propulsion, health management, long-life & high performance engine. These system level studies by a reusable flight demonstrator will be conducted for next three years.

Although many rockets have been launched so far, those conventional rockets have some problems such as high launching cost due to lack of reusability. Therefore, in Japan, the development of a reusable vertical-takeoff-and-vertical landing (VTVL) rocket vehicle is being promoted, and the nose entry system is adopted as a returning flight system, in which the attitude changes (=turnover) by aerodynamic force, the engine re-ignites, and then the rocket lands on the ground. In order to accomplish this turnover safely, it is known to be necessary to reduce the difference between the maximum value and the minimum value of C-m (the pitching-moment coefficient). In this research, we attached the delta wing with vortex flaps as fins to the reusable rockets in order to improve C-m characteristics during turnover. By attaching fins (the flap deflection angle is 0 degrees: Flap_0), the nose-up C-m becomes smaller than the case without fins (Body alone) at forward angles (angles of attack 0 - 90 degrees), but unfortunately, the nose-down C-m becomes larger at backward angles (AOA 90 - 180 degrees). On the other hand, by setting the flap deflection angle to -30 degrees (Flap_-30), the nose down C-m becomes smaller than that of Flap_0 at backward angles. Therefore, by setting Flap_0 at forward angles and Flap_-30 at backward angles, the difference between the maximum value and the minimum values of C-m can be reduced (12% smaller than Body-alone).

On February 3, 2018 at the JAXA Uchinoura Space Center, JAXA experimented SS-520 No. 5 launch with a 3U sized cube sat called TRICOM-1R aboard. After liftoff, flight of SS-520 No. 5 proceeded normally. Around 7 minutes 30 seconds into flight, TRICOM-1R separated and was inserted into its target orbit. And the launcher became the world's smallest class satellite launcher. SS-520 launch vehicle is one of sounding rockets operated in JAXA/ISAS, and originally two stage rocket. In this experiment, to make this vehicle put a satellite into orbit, the third stage motor is added. And this sounding rocket has four tail fins for spin stabilization, but usually don't have an attitude control system during the flight. But in this mission, it is needed to control its attitude to ignite second and third motor toward horizontal after first stage bum-out. The gas jet system is installed into between the first stage and the second stage of the vehicle as a unique active attitude control system. The gas jet system can control the spin axis direction and the spin rate of the vehicle during the coasting fight. Because of this constraint, the apogee altitude after the burn out of the first stage motor almost correspond with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket-based Nano launcher is planned to put TRICOM-1R into the elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. Because the perigee altitude is relatively low, the orbit life is very short. One of the mission requirements is to make the vehicle an orbit insertion with more than 30 days orbital lifetime. The vehicle error or the environment error deeply affect the achieved trajectory. These errors must be small enough to put TRICOM-1R into orbit. This paper discusses about the trajectory design on how to manage the sounding rocket into a satellite launching vehicle, the effect of the orbital distribution depending on the various errors, the flight safety analysis, and finally flight performance evaluation.

Copyright © 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. The development of a fully reusable vertical-takeoff, vertical-landing rocket is indispensable for reducing space transportation costs. However, there are many technical issues associated with such vehicles, such as the safe execution of a turnover maneuver during return flight. It is known that a relatively desirable pitching moment characteristic for turnover can be accomplished byemploying a slender-body configuration, but the reason for this is not well understood. In this study, a delayed detached-eddy simulation on the aerodynamic characteristics of such a slender-bodied reusable rocket is carried out for angles of attack between 0 and 180 deg using unstructured compressible computational fluid dynamics. Inviscid calculations are also conducted to distinguish the pitching moment contribution of the body configuration itself from the effects of viscosity and turbulence. It was found that two types of vortices were formedat0-90 deg, and these vortices affected the pitching moment distribution. Three types of vortices generated at 90-180 deg were also observed. Combined with the results of the inviscid simulation, it is concluded that the pitching moment characteristic is greatly impacted by the behaviors of these vortices and bubbles.

© 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. For the prediction of heat transfer coupled with sloshing phenomena in the propellant tanks of reusable launch vehicle for sounding mission, the pressure drop induced by heat transfer and the dynamic motion of liquid in sub-scale vessels were experimentally investigated. The correlation between the pressure drop and liquid motion was confirmed in the experiment. It was suggested that splash and wavy surface induced by violent motion of liquid cause the pressure drop in the closed vessel. In addition, as the preliminary investigation, non-isothermal sloshing of liquid nitrogen and liquid hydrogen were successfully visualized and pressure drop depending on the gaseous species was discussed.

© 2018 Univelt Inc. All rights reserved. On February 3, 2018 at the JAXA Uchinoura Space Center, JAXA experimented SS-520 No. 5 launch with a 3U sized cube sat called TRICOM-1R aboard. After liftoff, flight of SS-520 No. 5 proceeded normally. Around 7 minutes 30 seconds into flight, TRICOM-1R separated and was inserted into its target orbit. And the launcher became the world’s smallest class satellite launcher. SS-520 launch vehicle is one of sounding rockets operated in JAXA/ISAS, and originally two-stage rocket. In this experiment, to make this vehicle put a satellite into orbit, the third stage motor is added. And this sounding rocket has four tail fins for spin stabilization, but usually don’t have an attitude control system during the flight. But in this mission, it is needed to control its attitude to ignite second and third motor toward horizontal after first stage burn-out. The gas jet system is installed into between the first stage and the second stage of the vehicle as a unique active attitude control system. The gas jet system can control the spin axis direction and the spin rate of the vehicle during the coasting fight. Because of this constraint, the apogee altitude after the burn out of the first stage motor almost correspond with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket-based Nano launcher is planned to put TRICOM-1R into the elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. Because the perigee altitude is relatively low, the orbit life is very short. One of the mission requirements is to make the vehicle an orbit insertion with more than 30 days orbital lifetime. The vehicle error or the environment error deeply affect the achieved trajectory. These errors must be small enough to put TRICOM-1R into orbit. This paper discusses about the trajectory design on how to manage the sounding rocket into a satellite launching vehicle, the effect of the orbital distribution depending on the various errors, the flight safety analysis, and finally flight performance evaluation.

© 2018 Univelt Inc. All rights reserved. Although many rockets have been launched so far, those conventional rockets have some problems such as high launching cost due to lack of reusability. Therefore, in Japan, the development of a reusable vertical-takeoff-and-vertical-landing (VTVL) rocket vehicle is being promoted, and the nose entry system is adopted as a returning flight system, in which the attitude changes (=turnover) by aerodynamic force, the engine re-ignites, and then the rocket lands on the ground. In order to accomplish this turnover safely, it is known to be necessary to reduce the difference between the maximum value and the minimum value of Cm (the pitching-moment coefficient). In this research, we attached the delta-wing with vortex flaps as fins to the reusable rockets in order to improve Cm characteristics during turnover. By attaching fins (the flap deflection angle is 0°: Flap_0), the nose-up Cm becomes smaller than the case without fins (Body-alone) at forward angles (angles of attack 0 – 90 degrees), but unfortunately, the nose-down Cm becomes larger at backward angles (AOA 90 – 180 degrees). On the other hand, by setting the flap deflection angle to -30° (Flap_-30), the nose-down Cm becomes smaller than that of Flap_0 at backward angles. Therefore, by setting Flap_0 at forward angles and Flap_-30 at backward angles, the difference between the maximum value and the minimum values of Cm can be reduced (12% smaller than Body-alone).

© 2018 Univelt Inc. All rights reserved. In order to make the access to space for scientific researches much easier and make the opportunities of the rocket launches much frequent, a fully reusable sounding rocket is proposed in ISAS/JAXA. The mission definition of the proposed reusable sounding rocket are 1) To achieve 100km in altitude and returns to the launch site, 2) The 100kg payload to be carried, 3) Flight frequency is higher than 10 times per a year, 4) The minimum flight interval is one day, and 5) Operational flight cost should be an order of magnitude less than the existing ISAS sounding rocket. Reusable sounding rocket is different from the present expendable rockets in 1) repeated operations, 2) returning flight / reignition of engine / vertical landing, that is, 3) fault tolerant / health management. Some key technologies related to these characteristics of reusable system have been verified to design an operative reusable sounding rocket in phase-A. Technologies verifications respect to the reusable vehicle, 1) reusable engine development and repeated operations, 2) reusable insulation development for cryogenic tank, 3) aerodynamic design and model flight demonstration for returning flight, 4) cryogenic liquid propellant management demonstration, 5) landing gear development, and 6) health management system construction, have been successfully conducted from 2010 to 2016.After these technical demonstrations, we are proceeding with a study for system level verifications by a flight demonstrator from 2016 as the next step for the development of reusable sounding rocket. In this plan, a small test vehicle will be established for repeated flight demonstrations. Objectives of the demonstration are 1) system architecture study for repeated flight operation such as quick turnaround operation and fault tolerant design method, 2) life controlled and frequently repeated use of cryogenic propulsion system and its flight demonstrations, 3) study for the advanced returning flight method of vertical landers and its flight demonstrations, and 4) demonstration of advanced technology for future RLVs such as more composite on board, in flight fuel management, GH2/GOX auxiliary propulsion, health management, long-life & high performance engine. These system level studies by a reusable flight demonstrator will be conducted for next three years.

<p>In this paper, anomaly detection that is configured as a combination of state observer and Mahalanobis-Taguchi (MT) method is proposed for real time fault detection of rapid and dynamic phenomena such as rocket engine operation. Real time anomaly detecting is recognized as one of the most important elements to realize advanced reusable space transportation system. Conventionally, bottom-up type anomaly detecting logic based on failure mode and effect analysis (FMEA) is usually used for this purpose, however, it requires large amount of time and labor. The proposed method can improve this process. In the present method, error values between calculated ones through rocket engine simulator constructed on autoregressive moving average model and extended Kalman filter (EKF) and measured ones are standardized with existing normal operation data of the rocket engine so as to compute Mahalanobis' distance, which expresses degree of anomaly. We performed engine hot firing tests in simulated anomaly conditions. The obtained data was processed with the present method, and the simulated anomaly in the tests was detected as expected.</p>