野中 聡

The magneto-plasma sail (mini-magnetospheric plasma propulsion) produces the propulsive force due to the interaction between the artificial magnetic field around the spacecraft inflated by the plasma and the solar wind erupted from the Sun with a speed of 300-800 km/s. The principle of the magneto-plasma sail is based on the magnetic sail whose original concept requires a huge mechanical coil structure, which produces a large magnetic field to capture the energy of the solar wind. Meanwhile in the case of the magneto-plasma sail, the magnetic field will be expanded by the inertia of plasma flow to a few tens of kilometer in diameter, resulting in a thrust of a few Newton R. Winglee's group of the University of Washington originally proposed the idea of magnetic field inflation by the plasma. This paper investigates the characteristics of the magneto-plasma sail by comparing it with the other low-thrust propulsion systems (i.e., electric propulsion and solar sail), and the potential of its application to near future outer planet missions is studied. Furthermore, an engineering validation satellite concept is proposed in order to confirm the propulsion system specification and operation methodology. The main features are summarized as: (1) The satellite mass is around 180 kg assuming the H-IIA piggyback launch. (2) Since the magnetopause of the Earth magnetosphere is about 10 Re at Sun side and the bow shock is located at about 13 Re from the Earth, the satellite is injected into an orbit with 250 km perigee altitude and 20 Re apogee distance where apogee is located at the Sun side. (3) The magneto-plasma sail is turned on only in the vicinity of apogee outside the Earth's magnetosphere. (4) The thrust is estimated by the orbit determination result, and the plasma wind monitor is installed on the satellite to establish the relationship between the solar wind and the thrust. © 2005 Elsevier Ltd. All rights reserved.

A vertical take-off and vertical landing rocket is one of the future space transportation vehicle expected as fully reusable system. In the landing phase of vertical lander, the vehicle is decelerated by the main engine thrust and lands softly to the ground site. Then its aerodynamic characteristics are affected by the interaction between the engine plume and the subsonic free-stream against the vehicle. In order to investigate the influence of such interaction, wind tunnel tests were conducted. The aerodynamic forces and surface pressure were measured by using scale model of the Reusable Vehicle Testing (RVT) which is a small vehicle built for flight tests in ISAS/JAXA. Flowfield around the vehicle model was visualized by using Particle Image Velocimetry (PIV) method. As a result, the drag force and pitching moment acting the vehicle were affected by the change of pressure distribution due to the jet/free-stream interaction.

Computations of opposing jet flow from the vertical landing rocket vehicle are performed by using Large Eddy Simulation technique. Calculated results are compared with experimental data obtained by several wind tunnel testing. From the comparison of time-averaged flow features between calculation and PIV measurement, it is shown that dominant mean flow structures around the vehicle are nearly reproduced in this calculation. It is also shown that although a quantitative agreement of measured pressure coefficient values with those given by the present calculation is yet to be accomplished, the general trends in the measured values are reproduced well in the present calculation.

A baseline system of a reusable sounding rocket and its studies conducted in ISAS (Institute of Space and Astronautical Science) / JAXA (Japan Aerospace Exploration Agency) are presented. The vehicle adopts a vertical take-off and vertical landing (VTVL) system to minimize ground support equipments (GSE) and turn-around time. An integrated propulsion system which consists of four 2-ton-class liquid-hydrogen / liquid-oxygen (LH2/LOX) expander cycle engines, a gaseous-hydrogen / gaseous-oxygen (GH2/GOX) reaction control system (RCS) and an auxiliary power unit (APU) are introduced in order to simplify ground operations. The vehicle is designed to meet the requirements of carrying 100kg payload on a round trip up to 100km and repeating the mission within 24 hours. The airframe shape is basically axisymmetric; wings are excluded or minimum. The pros and cons of two ways of descent, i.e., nose-forward and base-forward descents, are discussed according to the vehicle system requirements; the nose-forward descent is adopted. The airframe shape is optimized so that the drag is minimum in ascent and the lift-drag ratio is maximum in descent. The 'turn-over' maneuver, i.e., the change from the nose-forward descent attitude to the base-down landing attitude, is required for the nose-forward descent system.

A fully reusable rocket vehicle is proposed to demonstrate good operability characteristics both on the ground and in flight. The proposed vehicle is to be used as a sounding rocket and has the capabilities of ballistic flight, returning to the launch site, and landing vertically making use of clustered liquid hydrogen rocket engines. Before initiating the development of this type of reusable rocket, a small test vehicle with a liquid hydrogen rocket engine was built and flight-tested. A demonstration of vertical landing and exercise of turnaround operation for repeated flights are the major objectives of the test vehicle. Three series of flight tests were performed in 1999, 2001 and 2003, and the flight test operation provided repeated flight environment and many valuable lessons were learned for designing the fully reusable rocket vehicle. © 2005 Published by Lister Science.

In October 2003, a new space agency, JAXA (Japan Aerospace Exploration Agency) was reorganized and started as a primary space agency to promote all space activities in Japan. The Institute of Space and Astronautical Science (ISAS) belonged to JAXA and continued to promote space science and technologies using unique scientific satellites, sounding rockets and balloons. This paper summarizes sounding rocket and ballooning activities of ISAS in the fiscal year of 2003 and 2004, associated with satellite launch programs. In this time period, three sounding rockets and nineteen balloons were launched by ISAS. One of the sounding rocket, S-310-35 was an international collaboration between Japan and Norway, which was launched from Andoya Rocket Range (ARR), Andenes, Norway, so as to study the upper atmospheric dynamics and energetics associated with the auroral energy in the polar lower thermosphere. Through the combination with the national researchers and the cooperation with international organizations, ISAS will keep its own flight opportunities and be able to obtain many new scientific findings.

For the prediction of sloshing in the propellant tank of a rocket vehicle, the preliminary investigation was conducted. The flow field in the propellant tank during the ballistic flight of the vehicle was experimentally reproduced with the sub-scale model. The lateral acceleration as large as about 0.8 G was provided with a mechanical exciter and the deformation of the liquid surface in the small vessel was visualized with a high-speed camera. The sloshing phenomena were also simulated with the CFD code, called CIP-LSM. The important features of surface deformation and wave breaking were successfully reproduced in the computation.

A baseline system of a sub-orbital reusable rocket and its aerodynamic studies conducted in ISAS (Institute of Space and Astronautical Science) / JAXA (Japan Aerospace Exploration Agency) are presented. The vehicle adopts a vertical take-off and vertical landing (VTVL) system to minimize ground support equipments (GSE) and turn-around time. An integrated propulsion system which consists of four 2-ton-class liquid-hydrogen / liquid-oxygen (LH2/LOX) expander cycle engines, a gaseous-hydrogen / gaseous-oxygen (GH2/GOX) reaction control system (RCS) and an auxiliary power unit (APU) are introduced in order to simplify ground operations and reduce turn-around time. The vehicle is designed to meet the requirements of carrying 100kg payload on a round trip to 120km and repeating the mission within 24 hours. Aerodynamic design considerations are made on the vehicle. The airframe shape is basically axisymmetric; wings are excluded or minimum. The pros and cons of two ways of descent, i.e., nose-forward and base-forward descents, are discussed according to the vehicle system requirements. Since in the base-forward descent the downrange requirement is not met, the nose-forward descent is adopted. The airframe shape is optimized combined with the aerodynamic force calculation so that the drag is minimum in ascent and the lift-drag ratio is maximum in descent. The 'turn-over' maneuver, i.e., the change from the nose-forward descent attitude to the base-down landing attitude, is required for the nose-forward descent system. The interaction between engine plumes and a free-stream become important for design of an attitude control system in the phase of deceleration using engines, and the interaction between an engine jet plume and a ground become important for design of a thermal protection system for landing gears just before the touch-down as well as the take-off. © 2004 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A fully reusable rocket vehicle is proposed to demonstrate good operability characteristics both on the ground and in flight. The proposed vehicle is to be used as a sounding rocket and has the capabilities of ballistic flight, returning to the launch site, and landing vertically making use of clustered liquid hydrogen rocket engines. Before initiating the development of this type of reusable rocket, a small test vehicle with a liquid hydrogen rocket engine was built and flight-tested. A demonstration of vertical landing and exercise of turnaround operation for repeated flights are the major objectives of the test vehicle. Three series of flight tests were performed in 1999, 2001 and 2003, and the flight test operation provided repeated flight environment and many lessons were learned valuable for designing the fully reusable rocket vehicle.

For achieving a fully reusable rocket vertical as a future space transportation system, the conceptual designs of vehicle systems and flight tests by a small test vehicle are presently being conducted in ISAS/JAXA. In this system design, aerodynamic design considerations are made on a vertical take-off and vertical landing vehicle. One of the considerable issues of a vertical lander is the effect of the interaction between a supersonic nozzle jet and a free-stream when the vehicle is decelerated by the main engine thrust in the landing phase. In order to investigate the influence of such counter-flow interaction in detail, wind tunnel tests were conducted in low speed wind tunnel in ISAS and ISTA/JAXA. The aerodynamic forces and pressure on the base surface were measured by using a scale model of the vehicle. The flowfield around the model was visualized by using smoke and tuft. The velocity distribution was measured by a particle image velocimetry (PIV) technique. The aerodynamic characteristics in the vertical landing phase are affected by not only the reduction of the base pressure but also the non-separated flow around the model side.

The magneto-plasma sail (mini-magnetospheric plasma propulsion) produces the propulsive force due to the interaction between the artificial magnetic field around the spacecraft inflated by the plasma and the solar wind erupted from the Sun with a speed of 300-800 km/s. The principle of the magneto-plasma sail is based on the magnetic sail whose original concept requires a huge mechanical coil structure, which produces a large magnetic field to capture the energy of the solar wind. Meanwhile in the case of the magneto-plasma sail, the magnetic field will be expanded by the inertia of plasma flow to a few tens of km in diameter, resulting in a thrust of a few N. R.Winglee's group of the University of Washington originally proposed the idea of magnetic field inflation by the plasma. This paper investigates the characteristics of the magneto-plasma sail by comparing it with the other low-thrust propulsion systems (i.e., electric propulsion and solar sail), and the potential of its application to near future outer planet missions is studied. Furthermore, an engineering validation satellite concept is proposed in order to confirm the propulsion system specification and operation methodology. The main features are summarized as: The satellite mass is around 180kg assuming the H-IIA piggyback launch. 2) Since the magnetopause of the Earth magnetosphere is about 10Re at Sun side and the bow shock is located at about 13Re from the Earth, the satellite is injected into an orbit with 250km perigee altitude and 20 Re apogee distance where apogee is located at the Sun side. 3) The magneto-plasma sail is turned on only in the vicinity of apogee outside the Earth's magnetosphere. 4) The thrust is estimated by the orbit determination result, and the plasma wind monitor is installed on the satellite to establish the relationship between the solar wind and the thrust.

資料番号: SA0200020000

Shock shapes over a sharp cone in the intermediate hypersonic flow regime are calculated to examine the validity of the existing two-temperature thermochemical model. Two different apex angles are considered in the calculations: one with a half-angle of 30 deg and the other with 45 deg. The calculations for these geometries are carried out for several different static pressure values at the flight velocity of about 3.0 km/s. The calculated shock layer thickness is compared with the corresponding experimental data obtained in a ballistic range. The results show that the two-temperature model well reproduces the experimental data for the flow conditions in which chemical reactions as well as vibrational excitations are absent. However, the calculated shock layer thickness tends to be thinner than the experimental data for the flow conditions in which vibrational excitation begins to occur. It is implied that vibrational relaxation has a close connection with the thinner shock layer. The study confirms our previous results that the shock layer thickness over a sphere in the same velocity range can be underestimated in the calculation using the existing two-temperature model.

For the prediction of sloshing in the propellant tank of rocket vehicle utilized in RVT (reusable rocket vehicle testing) conducted by ISAS, the flow field in the propellant tank during the coasting flight was experimentally reproduced with the sub-scale model of it. The lateral acceleration as large as about 1.0 G was provided with a mechanical exciter and the deformation of liquid surface in the vessel was visualized with a high-speed camera. The several configurations of damping devices were installed in the vessel, which should keep the ullage gas away from the outlet port, and tested to verify their performance. It was consequently suggested that the combination of a baffle plate and a perforated cylinder could be effective against the gas suction before the re-ignition of the engine. The sloshing was also simulated with the newly developed CFD code, called CIP-LSM. The numerical results showed good agreement with the corresponding data obtained in the experiment. For the appropriate assessment of liquid behavior in the flight of RVT, the flow field in the full scale tank was investigated numerically under the practical condition of acceleration. © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

If a dense plasma were exhausted near the center of a magneetic sail, the magnetic field could be expanded far away from the spacecraft, thus the energy of the solar wind can be captured by this huge magnetic field in spite of very low-density solar wind. Then the magnetic sail can propel a spacecraft by the solar wind in the inerplanetary space. Such a magnetoplasma sail was analytically studied, and large thrust to power ratio as much as 250mN/kW was explained. When applied to short-term deep space missions, the magnetoplasma sail has great advantage against other electric propulsion systems because of its ability to achieve larger thrust to power ratio. © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Some of solid rocket motor plumes are reported to cause a telecommunication black-out. To clarify the mechanism of this microwave-plume interaction, attenuations of the microwave signals were measured in three ground firing tests for Japanese solid rocket motors, M25, M14, and SRB-A. As a microwave diagnostics, multifrequency microwave technique was employed (S-band, 2.4 GHz, C-Band, 5.6 GHz, and X-Band, 8.4GHz), by which both the electron density and the electron collision frequency were simultaneously determined using theoretical attenuation by a plasma slab model as 3x1016 m-3 and 7x1010 Hz during the effective firing period of the motors. Although this successful determination of plume plasma properties indicated the cause of the telecommunication black-out is plasma, however, near the end of the firing, large attenuations and departure from the theoretical curve will imply another possible attenuation mechanism in addition to the plasma effect depending on the firing period. © 2002 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

For achieving fully reusable rocket vehicle, aerodynamic design considerations are made on vertical landing vehicle. The vehicle body is based on an axisymmetric conical and cylindrical geometry. Two types of entry flight concepts for the vertical landers, nose entry and base entry, are discussed. The relations between vehicle shapes and its aerodynamic characteristics are investigated by trade-off studies about each entry concepts. Entry flight capabilities of these vehicles are evaluated, and considerable problems in respect to aerodynamics are pointed out Aerodynamic issues in landing phase at low altitude and low flight speed are also discussed. The possibilities for realization of VTOL SSTO rocket are synthetically evaluated from aerodynamic point of view. A base entry vehicle with low fineness ratio has characteristics of small downrange, low heat load, and large ascent drag loss. On the other hand, characteristics for nose entry in respect to downrange and drag loss is advantageous. Characteristics of a base entry sounding rocket proposed in ISAS is discussed. © 2001 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The theoretical analysis describing the dependence of the shock standoff distance on a sphere was modified for application to cases where vibrational relaxation behind the bow shock is present and chemical dissociation effects are negligible. The analysis demonstrates which flow parameters are most suitable for the correlation of experimental data and allow standoff distances to be estimated quickly without the need to run time computational fluid dynamics codes.

Hypervelocity impacts produce extreme pressures in both the projectile and target materials. In the present study the hypervelocity impact crater formation is investigated experimentally both from basic gasdynamic and metallurgical view points. The debris cloud generated by the impact of a 14 mm diameter high density polyethylene projectile flying at 4.0 km /s on a 2 mm thick aluminium alloy (2024) shield is visualized by shadowgraphy. The debris cloud formation is also simulated numerically using AUTODYNE ™ - 2Dhydrocode. Experimentally visualized debris cloud shape agrees well with numerical simulation results. Although the average value of Vicker's micro-hardness for AI 2024 is 138 Kgf/mm2, in the vicinity of the edge of the impact crater the value increased to ~ 170 Kgf/mm2. The increase in hardness due to strain hardening seems to be off set by localized adiabatic heating. Spallation and plugging rings are observed even on the back side of the target. Adiabatic shear bands which nucleate the cracks for spading, non- homogenous secondary phases of plastic deformation appear to be the possible mechanism of fracture in hypervelocity impacts. © 2000 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The density distribution in shock layer over a sphere is measured in a ballistic range by using finite fringe holographic interferometry. The measurement is made for a hemisphere of nose-radius of 15 mm, flight speed of 2.53 km/s, and initial pressure of 5,065 Pa. The obtained finite fringe interferogram is analyzed by Fourier transform techniques and an Abel deconvolution method to produce density maps. From the obtained density distribution, the density increase due to the real-gas effects was clearly recognized. The density profiles are compared with nonequilibrium calculation by using two-temperature model, and the numerical results well reproduce the density profiles. © 2000 by American Institute of Aeronautics and Astronautics, Inc. All rights reserved.