Daisuke Akita

Kinetic Analysis on Plasma Flow of Solar Wind Around Magnetic Sail

The magnetic sail utilizes the momentum of the solar wind (high-speed plasma flow) to produce the thrust by using an artificial magnetic field. Since the momentum flux from the sun is transformed into the thrust like a solar sail, no propellant is required and infinite Isp performance is available. Hence, it has a potential to reduce drastically the duration for a deep space mission, since it can accelerate the spacecraft up to its theoretical limit, that is, the speed of solar wind. In the present study, a small magnetic sail of 10m-size, which seems to be reasonable size as a test vehicle realizable within the present technology level on the space structure, is considered. Its acceleration performance and scales of the phenomena in relation to the solar wind flow around the magnetic sail are roughly estimated. Based on the estimation, the interaction of the solar wind with the applied magnetic field is numerically simulated by the full particle (PIC) method. Fundamental features of the flow field and the induced electromagnetic field around the small magnetic sail are clarified. Force acting upon the magnetic sail is also estimated by considering the Lorentz force generated by the induced electromagnetic field and the momentum change of the solar wind around the magnetic sail.

Deployment and Flight Test of Inflatable Membrane Aeroshell using Large Scientific Balloon

A flexible aeroshell for atmospheric entry vehicles has attracted attention as an innovative space transportation system because the aerodynamic heating during an atmospheric entry can be reduced dramatically due to its low ballistic coefficient. We have researched and developed a capsule-type vehicle with a flare-type membrane aeroshell sustained by an inflatable torus frame. One of the key technologies is to develop a large and low-mass aeroshell including inflatable structures. As a part of the development, a miniature experimental vehicle was developed and a balloon drop test was carried out in order to acquire the vehicle with inflatable structures in a high altitude and a free flight condition. The diameter, the total mass, and the ballistic coefficient of the experimental vehicle are 1.264m, 3.375kg, and 2.69kg/m 2 , respectively and its aeroshell consists of a thin membrane flare made of nylon and a torus which can be inflated by gas pressure. The inflatable aeroshell was deployed and the experimental vehicle was separated from the balloon at an altitude of 25km. After the separation, the vehicle flied 30 minutes until a splashdown. This balloon test is very successful and fruitful and following results were achieved. 1) Remote deployment system of the inflatable aeroshell by sending a command from a ground station

Reentry Demonstration of Flare-type Membrane Aeroshell for Atmospheric Entry Vehicle using a Sounding Rocket

A flexible aeroshell for atmospheric entry vehicles has attracted attention as an innovative space transportation system because the aerodynamic heating during the atmospheric entry can be reduced dramatically due to its low ballistic coefficient. We carried out wind tunnel tests, numerical simulations and flight demonstrations using balloons, focusing on a capsule-type vehicle with a flare-type membrane aeroshell sustained by an inflatable torus frame. For the next step, a reentry demonstration using a sounding rocket is planned as a important milestone. The experimental vehicle which has a 1.2-meter diameter flare-type thin membrane aeroshell sustained by an inflatable torus is being developed for the reentry demonstration. In the reentry demonstration using a sounding rocket, the experimental vehicle reenters the atmosphere from an altitude of 150 km and experiences a hypersonic free flight where the Mach Number is 4.5 and the moderate aerodynamic heating where the heat flux at a stagnation is about 20kW/m. The flight trajectory, the behavior of the aeroshell and the aerodynamic heating condition will be measured by onboard sensors and a telemetry system during the reentry. The performance of the flexible aeroshell as a decelerator for reentry vehicles can be demonstrated in this flight test and the results and knowledge obtained in ground tests beforehand can be validated using this flight data.

Development of flare-type inflatable membrane aeroshell for reentry demonstration from LEO

An inflatable decelerator is promising for a next generation atmospheric-entry system, because it can be packed compactly in the launch and cruise phase and it can be deployed to a large aerodynamic device in the atmospheric-entry phase. Our group has researched and developed this technology since 2000, focusing on a flare-type membrane aeroshell sustained by a single inflatable ring, especially. In our activity, the re-entry demonstration using a Japanese S-310 sounding rocket was carried out successfully in 2012. As a next millstone of our research and development, the re-entry demonstration from the low earth orbit is planned utilizing an opportunity for piggy-back satellites. The overview of the planned reentry demonstration is introduced in this paper. There are several important technical issues to overcome in order to realize this demonstration. Two important issues of these is also introduced. First topic is the structural strength tests using a low-speed wind tunnel to understand the structural strength of a large flare-type membrane aeroshell supported by a single inflatable ring. Second topic is an evaluation on the thermal durability of inflatable structures using a newly developed inductively coupled plasma heater.