山田 和彦

A numerical model for simulating air and nitrogen inductively coupled plasmas (ICPs) was developed considering thermochemical nonequilibrium and the third-order electron transport properties. A modified far-field electromagnetic model was introduced and tightly coupled with the flow field equations to describe the Joule heating and inductive discharge phenomena. In total, 11 species and 49 chemical reactions of air, which include 5 species and 8 chemical reactions of nitrogen, were employed to model the chemical reaction process. The internal energy transfers among translational, vibrational, rotational, and electronic energy modes of chemical species were taken into account to study thermal nonequilibrium effects. The low-Reynolds number Abe-Kondoh-Nagano k-epsilon turbulence model was employed to consider the turbulent heat transfer. In this study, the fundamental characteristics of an ICP flow, such as the weak ionization, high temperature but low velocity in the torch, and wide area of the plasma plume, were reproduced by the developed numerical model. The flow field differences between the air and nitrogen ICP flows inside the 10-kW ICP wind tunnel were made clear. The interactions between the electromagnetic and flow fields were also revealed for an inductive discharge. Published by AIP Publishing.

<p>For landing a rover on the Mars ground, supersonic parachute has been developed in JAXA. Key technologies are categorized in aerodynamic performance, mechanical strength, ejection system, and validation method of the design for pre-flight model. So far, we have performed experiments in low-speed, transonic, and supersonic wind tunnels in Chofu aerospace center and ISAS. From these experiments, we have investigated aerodynamic performance such as drag coefficients, opening load factor, and stability of the parachute. We have also evaluated the mechanical strength in these wind tunnel tests. In addition, ejection system with automobile airbag inflator has been developed and a vertical ground test is performed in Noshiro Rocket Testing Center.</p>

A demonstration flight of an advanced reentry vehicle was carried out using a sounding rocket. The vehicle was equipped with a flexible (membrane) aeroshell deployed by an inflatable torus structure. Its most remarkable feature was the low ballistic coefficient that enables reduction in aerodynamic heating and deceleration at a high altitude. During the suborbital reentry, temperatures at several locations on a backside of the flexible aeroshell and inside the capsule were measured by means of embedded thermocouples. The aerodynamic heating behavior of the vehicle was investigated using the measured temperature history, in combination with a numerical prediction in which a flow-field simulation of the heating was conducted. In this flow-field simulation, both laminar flow and turbulent flow were assumed, and the deformation of the flexible aeroshell was considered. A thermal model of the capsule and membrane aeroshell was developed, and the heat flux profiles of the vehicle surface during aerodynamic heating were constructed based on the measured temperatures. The measured temperature data were found to be in reasonable agreement with the predicted data if the flow field near the capsule of the vehicle was assumed to be laminar, with a transition to turbulent flow near the membrane aeroshell.

Two-dimensional (2D) numerical simulations of thermochemical nonequilibrium inductively coupled plasma (ICP) flows inside a 10-kW inductively coupled plasma wind tunnel (ICPWT) were carried out with nitrogen as the working gas. Compressible axisymmetric NavierStokes (N-S) equations coupled with magnetic vector potential equations were solved. A fourtemperature model including an improved electron-vibration relaxation time was used to model the internal energy exchange between electron and heavy particles. The third-order accuracy electron transport properties (3rd AETP) were applied to the simulations. A hybrid chemical kinetic model was adopted to model the chemical nonequilibrium process. The flow characteristics such as thermal nonequilibrium, inductive discharge, effects of Lorentz force were made clear through the present study. It was clarified that the thermal nonequilibrium model played an important role in properly predicting the temperature field. The prediction accuracy can be improved by applying the 3rd AETP to the simulation for this ICPWT.

An inflatable decelerator is promising as a next-generation atmospheric-entry system owing to its reduced aerodynamic heating and high packing efficiency. In this study, a suborbital reentry demonstration of a flare-type thin-membrane aeroshell sustained by a single inflatable torus using an S-310-41 sounding rocket was carried out. An experimental vehicle was specially developed for this reentry demonstration; it was equipped with a 1.2-m-diam flare-type thin-membrane aeroshell and had a total mass of 15.6 kg. In the flight test, the aeroshell with an inflatable torus was deployed at an altitude of 100 km during a suborbital flight under the conditions of zero-gravity and near vacuum. The experimental vehicle reentered Earth's atmosphere from an altitude of 150 km. During free fall, it accelerated to a Mach number of 4.5 (1.32 km/s) because of gravity force. After that, it started decelerating because of aerodynamic force at an altitude of 70 km. According to the flight data, the vehicle remained intact during the reentry and the aeroshell achieved the expected decelerating performance. This reentry demonstration proves that the flare-type thin-membrane aeroshell sustained by the inflatable torus works well as a decelerator for atmospheric-entry vehicles. Further, the drag coefficient of the experimental vehicle in the supersonic, transonic, and subsonic regimes under free-flight conditions was estimated from the flight trajectory.

An inflatable re-entry vehicle is a candidate for future re-entry systems. Owing to the large area and configuration of the vehicle, it can afford a few advantages during the re-entry, descent, and landing approach, such as a decrease of aerodynamic heating and soft landing without requiring a parachute system. To investigate aerodynamic characteristics of inflatable reentry vehicle at low-Mach-number flight, wind tunnel tests were performed in JAXA Low-Speed-Wind tunnel. In this research, we investigated aerodynamic characteristics of 2 types of inflatable reentry vehicle, SMAAC and TITANS, at a low-Mach-number by using numerical simulation. Through the flow field simulation, it was indicated that the computed result of drag coefficient shows reasonable agreement with the experimental one. In the case of TITANS, the computed results showed good agreements compared with experimental results though it was confirmed that a blockage effect was observed.

A numerical simulation model that combines the plasma flows and electromagnetic waves around a reentry vehicle during atmospheric reentry was developed to evaluate the radio frequency blackout and plasma attenuation. The physical properties of the plasma flow in the shock layer and wake region were obtained using a computational fluid dynamics technique. The electromagnetic waves were expressed using a frequency-dependent finite difference time domain method with the plasma properties. Combined simulations were performed for the atmospheric reentry demonstrator of the ESA at various altitudes based on reentry orbit data. The electromagnetic wave behaviors around the vehicle during atmospheric reentry were investigated in detail. Moreover, a parametric analysis with different ionization reaction models was performed. It was confirmed that the vehicle is surrounded by the plasma and the propagation of the electromagnetic waves is prevented at high altitude. Then, the plasma is dissipated and the propagation recovers at low altitude. Validation of the simulation model was performed based on the plasma attenuation history of the experimental flight data. A comparison of the measured and predicted results showed good agreement. It was concluded that the combined simulation model could be an effective tool for investigating the radio frequency blackout and the plasma attenuation of radio wave communication.

Numerical simulations of 10 kW and 110 kW inductively coupled plasma (ICP) wind tunnels were carried out to study physical properties of the flow inside the ICP torch and vacuum chamber with air as the working gas. Two-dimensional compressible axisymmetric NavierStokes (N-S) equations that took into account 11 species and 49 chemical reactions of air, were solved. A heat source model was used to describe the heating phenomenon instead of solving the electromagnetic equations. In the vacuum chamber, a four-temperature model was coupled with N-S equations. Numerical results for the 10 kW ICP wind tunnel are presented and discussed in detail as a representative case. It was found that the plasma flow in the vacuum chamber tended to be in local thermochemical equilibrium. To study the influence of operation conditions on the flow field, simulations were carried out for different chamber pressures and/or input powers. The computational results for the above two ICP wind tunnels were compared with corresponding experimental data. The computational and experimental results agree well, therefore the flow fields of ICP wind tunnels can be clearly understood.

Numerical simulations of the plasma flow and electromagnetic wave around a membrane-aeroshell type reentry vehicle were performed using various physical model combinations, and the possibility of radio frequency blackout of transceiver antenna embedded at the rear of the vehicle was investigated. The flowfield was assumed to be in thermochemical nonequilibrium, and it was described by the Navier-Stokes equations with a multitemperature model and the equation of state. The simulations were performed for several altitudes, including the highest heat flux point according to reentry orbit data. Through these computations, the detailed distributions of the flowfield properties in the shock layer and wake region were successfully obtained. To evaluate the possibility of radio frequency blackout during atmospheric reentry, the distribution of the electron number density around the inflatable vehicle was clarified. A frequency-dependent finite-difference time-domain method was used for simulations of electromagnetic waves, and the physical properties were obtained from the computational results of the plasma flow calculation. Electromagnetic wave behaviors in an ionized gas region behind the inflatable vehicle were investigated. It was found that the number density of electrons was sufficiently small and that the electromagnetic waves can propagate with no reflection and less attenuation. These results suggest that radio frequency blackout may not occur during the atmospheric reentry of the inflatable vehicle.

Numerical investigation of nonequilibrium inductively coupled plasma (ICP) flow was carried out to study the physical properties of the flow inside a 10-kW ICP torch with the working gas being nitrogen. The flow field was described by two-dimensional compressible axisymmetric Navier-Stokes (N-S) equations that took into account 5 species and 8 chemical reactions. The magnetic vector-potential equations were tightly coupled with the flow-field equations to describe the heating process by inductive discharge. A four-temperature model was adopted to model thermal nonequilibrium process in the discharge torch. The characteristics of ICP flow such as thermal nonequilibrium, inductive discharge, and strong effects of Lorentz forces became clear through the present study.

The electrodynamic effect on the partially ionized flow around magnetized bodies was numerically investigated. In particular, the double-cone model was considered because, unlike a simple blunt-nosed model, it generates a complex flow including shock-shock and shock-boundary-layer interactions. Such flows are significantly affected by an applied magnetic field. The modification of the flowfield due to the applied magnetic field causes drag force enhancement and/or the mitigation of the aerodynamic heating. This is enhanced with increasing magnetic field intensity. Furthermore, local flow features such as the separation bubble and the local peak heating that appears near the kink point of the double-cone model are significantly affected. The size of the separation bubble increases with increasing magnetic field intensity and is clearly influenced by the configuration of the magnetic field. The local peak heating is reduced with increasing magnetic field intensity, but the effect of the magnetic field configuration is weak.

An inflatable decelerator is promising as atmospheric entry systems in the next generation. Although the various kinds of inflatable decelerators were proposed and researched in the past, we focus especially on a flare-type membrane aeroshell sustained by an inflatable torus. The flare-type membrane aeroshell may easily make re-entry systems simple and light-weighted because the aeroshell can be composed mainly of light-weighted membrane. As an important milestone of the development of the vehicle, a re-entry flight demonstration of the vehicle with the flare-type membrane aeroshell was carried out using Japanese S-310 sounding rocket in August, 2012. This flight experiment was successfully accomplished and various data, including the attitude of the vehicle during the supersonic atmospheric re-entry, was acquired. The attitude dynamics was qualitatively reconstructed base on the flight data, which shows that the vehicle entered into the atmosphere with high angle of attack and experienced a drastic attitude motion caused by the aerodynamic force. After the initial significant motion phase, the attitude behavior of the vehicle calmed down before the successive deceleration phase. That is, the vehicle made a flight with a normal attitude and, as planned, its attitude was almost stable after the atmospheric-entry. In all, it was confirmed that the attitude stability of vehicles with a flare-type membrane aeroshell could be realized on a supersonic atmospheric-entry. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

An inflatable decelerator is promising as atmospheric-entry systems in the next generation thanks to the aerodynamic heating relaxation and its packing efficiency. Our group focuses on a flare-type membrane aeroshell sustained by an inflatable torus, especially. As an important milestone of our development, a re-entry demonstration of the flare-type membrane aeroshell was carried out using a Japanese S-310 sounding rocket. The experimental vehicle which has a 1.2-meter-diameter membrane aeroshell and 15.6kg in total weight was developed for the re-entry demonstration. In this flight test, the membrane aeroshell with the inflatable torus was deployed at 100km in altitude during a suborbital flight under the zero gravity and vacuum condition, and the experimental vehicle re-entered the earth atmosphere from 150km in altitude. The experimental vehicle accelerated to 1.32km/s and Mach Number 4.5 due to the gravity force and started decelerating due to the aerodynamic force at 70km in altitude. According to the flight data, the experimental vehicle kept intact during the re-entry and the flare type membrane aeroshell achieved the expected decelerating performance. This re-entry demonstration proves that the flare-type membrane aeroshell sustained by the inflatable torus works well as a decelerator for atmospheric-entry vehicles. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserve.

A reentry flight demonstration of an advanced reentry vehicle was carried out using a sounding rocket. The vehicle is equipped with a flexible (membrane) aeroshell deployed by an inflatable torus structure. During the reentry flight, temperatures at several locations of a backside of a flexible aeroshell and a sidewall of a capsule were measured by means of embedded thermocouples. Behavior of aerodynamic heating of the vehicle is investigated along with the measured temperature history, combining the numerical prediction in which the numerical flow simulation and the heating estimation based on it were conducted. In the flow field simulation, not only laminar flow but also turbulent flow were assumed, and modification of the flexible aero-shell were considered. It was found that the measured temperature data shows a reasonable agreement with the predicted one if the flow field near the membrane aeroshell of the vehicle is assumed to be turbulent flow. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A space vehicle with a large-area, low-mass membrane aeroshell has potential as a reentry system in the near future, because the large-area, low-mass membrane aeroshell dramatically reduces aerodynamic heating and achieves a soft landing without a conventional parachute system. To demonstrate membrane aeroshell technology, a drop flight test of a capsule-type experimental flight vehicle with a 1.5-m-diam, flare-type, flexible, deployable membrane aeroshell was carried out using a large scientific balloon. In this flight test, the experimental flight vehicle was dropped from the balloon at an altitude of 39 km and underwent free flight. The flight data collected using onboard sensors were transmitted successfully during the flight by the telemetry system. The data showed that the vehicle was stable in free-flight condition and that its flight path and aerodynamic characteristics agreed well with results from previous trajectory analysis and wind-tunnel tests. This test flight clearly demonstrated that this flare-type flexible aeroshell successfully functions as a stable decelerating device in free flight.