Kazutaka Nishiyama

Powered flight of electron cyclotron resonance ion engines on Hayabusa explorer

The electron cyclotron resonance ion engine has long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10s, each generating a thrust of 8 mN, specific impulse of 3200 s, and consuming 350 W of electric power, propelled the Hayabusa asteroid explorer launched on May 2003. After vacuum exposure and several baking runs to reduce residual gas, the ion engine system established continuous acceleration. Electric propelled delta-V Earth gravity assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, Hayabusa, using solar electric propulsion, managed to successfully cover the solar distance between 0.86 and 1.7 AU. It rendezvoused with, landed on, and lifted off from the asteroid Itokawa. During the 2-year flight, the ion engine system generated a delta-V of 1400 m/s while consuming 22 kg of xenon propellant and operating for 25,800 h.

Verification Tests of Carbon-Carbon Composite Grids for Microwave Discharge Ion Thruster

Anionbeam opticsfora10-cm-diam 400-W-classmicrowavedischargeion thrusterwasfabricatedanditsapplicabilityto along-termspacemissionwasdemonstrated.Theopticsconsistsofthree1-mm-thick e atcarbon ‐carbon composite panels with approximately 800 holes that were mechanically drilled and positioned with § 0:02-mmaccuracy.Whenmounted onanaluminum ring,spacingforthethreegridswaskeptat0.5 mm bythreesetsofspacers. The thruster produced an ion beam current of 140 mA with a microwave power of 32 W for plasma generation and a total acceleration voltageof 1.8 kV. Although thegrid is sputtered by the impingement of slow ions produced in charge-exchange collisions between fast beam ions and neutral atoms leaking from the engine, the grid showed only slight damage even after an 18,000-h endurance test. Also, other qualie cation tests including a mechanical test under launch conditions as well as a thermal vacuum test simulating the spacecraft thermal environment were successfully completed. Hence, the grid system was qualie ed for spacecraft propulsion.

Powered Flight of HAYABUSA in Deep Space

The electron cyclotron resonance ion engine, “μ10,” has a long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10, each generating a thrust of 8 mN, specific impulse of 3,200 seconds, and consuming 350 W of electric power, propel the “HAYABUSA” asteroid explorer that was launched on May 2003. After vacuum exposure and several runs of baking to reduce residual gas, the ion engine system established continuous acceleration. Delta-V Earth Gravity Assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, HAYABUSA, using solar electric propulsion, managed to successfully cover the distance between 0.86 AU and 1.7 AU in the solar system, as well as rendezvous with, land on, and lift off from the asteroid Itokawa. During the 3-year flight, the ion engine system generated a delta-V of 1,400 m/s while consuming 22 kg of xenon propellant and operating for 25,900 hours.

Hayabusa's Way Back to Earth by Microwave Discharge Ion Engines

The cathode-less electron cyclotron resonance ion engines, μ10, propelled the Hayabusa asteroid explorer, launched in May 2003, which is focused on demonstrating the technology needed for a sample return from an asteroid, using electric propulsion, optical navigation, material sampling in a zero gravity field, and direct re-entry from a heliocentric orbit. It rendezvoused with the asteroid Itokawa after a two year deep space flight with a delta-V of 1.4 km/s, 22 kg of xenon propellant consumption and 25800 hours of total accumulated operational time of all the four ion engines added up. Though it succeeded in landing on the asteroid on November 2005, the spacecraft was seriously damaged. This delayed the Earth return in 2010 from the original plan in 2007. Reconstruction on the operational scheme using remaining functions and newly uploaded control logic made Hayabusa leave for Earth in April 2007. After a coasting period of 2008, the ion propulsion was reignited in February 2009. Although most of the neutralizers were degraded and unable to be used by fall of 2009, a combination of an ion source and its neighboring neutralizer has been successfully operated for the last 3230 hours including a series of final trajectory correction maneuvers. Before reentry, the total accumulated operational time reached 39637 hours consuming a total of 47 kg Xenon propellant. Total duration of powered spaceflight is 25590 hours which provided a delta-V of 2.2 km/s and a total impulse of 1 MN·s, approximately. Finally, the spacecraft returned to Earth. Its reentry capsule, which may contain samples from asteroid Itokawa, was retrieved from the Australian outback according to plan .

Status of Microwave Discharge Ion Engines on Hayabusa Spacecraft

[Abstract] The μ10 cathode-less electron cyclotron resonance ion engines made the Hayabusa spacecraft rendezvous with the asteroid Itokawa in 2005. Though the spacecraft was seriously damaged after the successful soft-landing and lift-off, the xenon cold gas jets from the ion engines rescued the Hayabusa. New attitude stabilization method using a single reaction wheel, the ion beam jets, and the solar pressure was established and enabled the homeward journey aiming the Earth return on 2010. The total accumulated operational time of the ion engines reaches 28,000 hours at the end of May 2007.

Thrust Enhancement of a Microwave Ion Thruster

Two optical fiber measurement techniques are used in this paper to reveal the physical mechanism of the enhancement of the thrust force of the μ10 electron cyclotron resonance ion thruster. The beam current of the μ10 thruster was increased in previous studies by changing the propellant injection method. In this study, to observe the difference in plasma distributions, optical fiber probes were inserted into the thruster under beam acceleration. The first measurement was laser absorption spectroscopy. By traversing the optical fiber, the number densities of Xe I 5p5(P3/202)6s[3/2]20 1 were obtained along the center axis. The second measurement was an electric-optic element probe measurement conducted to measure the intensities of the microwave electric field. Both measurements suggest that there is plasma in the waveguide in the conventional model of the thruster. This phenomenon is possibly caused by the leakage of electrons from the electron cyclotron resonance region to the waveguide. As a result, this...

Ground Chamber Measurements of the Electromagnetic Emissions from the Hayabusa Ion Engine

DOI: 10.2514/1.19473 Radiated electric field emissions from the prototype model of the ion engine system of the asteroid explorer Hayabusa (MUSES-C) were measured in approximate accordance to MIL-STD-461C. The typical noise level exceeded the narrowband specification at frequencies less than 5 MHz. The microwave discharge neutralizer generates broadband noise and narrowband oscillations that have a fundamental frequency of about 160 kHz and are accompanied by its harmonics up to the fifth. Leakage of 4.25 GHz microwaves for plasma production and its second harmonic were 65 dB and 35 dB above specifications, respectively. The X-band receiver onboard Hayabusa measured the noise from the ion engine system at the uplink frequency of 7.16 GHz through a horn antenna. This susceptibility test showed that the microwave discharge ion thruster is unlikely to interfere with deep space microwave communication. Nomenclature d = neutralizer orifice diameter ID = electron emission current of the neutralizer l = anode distance from the neutralizer _ m = xenon flow rate P� = microwave power for plasma discharge of the neutralizer VD = voltage between the neutralizer and the anode in diode mode operation

Electron Cyclotron Resonance Ion Source for Ion Thruster

An electron cyclotron resonance (ECR) ion source was designed and its performance was studied in comparison with a DC ion source. Ion production cost of 340 V/ion for ECR ion source and 189 V/ion for DC ion source were obtained at the pressure of 0.5 mTorr and the discharge power of 50 W. A luminosity distribution of the Ar-II line was imaged through an interferometer filter and a luminous arch hill was observed near the ECR region, where the plasma was resonantly generated. Ion wall loss measurement showed that the ion production cost became worse due to the excessive ion wall loss near the cusped magnets, where the plasma confinement was inferior to that of the DC ion source.

Experimental Study for Enhancement Thrust Force of the ECR Ion Thruster μ10

In order to improve the thrust force of the ECR Ion Thruster μ10, two kinds of experiments were conducted. At the first experiment, a quartz plate was set at a waveguide to prevent the plasma from situating in the waveguide. Though it was introduced to improve the transmittance of microwave from the waveguide to a discharge chamber, the beam current did not increase. Secondly, a design of a spacer which locates between magnet rings was changed in two ways, different thick spacers and an anode spacer to improve the ion production cost. In addition to a 7-mm spacer (original) 4, 10, and 13-mm spacers were manufactured to eliminate the possibility of producing ions which cannot be accelerated by a screen grid. The anode spacer had a role to decrease ions which collide with the spacer. As a result, the higher spacer did not improve the beam current, however, the anode spacer improve the beam current by 5%.

Electric Field Measurement of ECR Ion Thruster u 10 with Optical Fiber Sensor

In order to reveal the internal phenomenon theoretically within the electron cyclotron resonance ion thruster μ 10, internal microwave electric field measurement is very important because it is closely related to plasma producing mechanism. We have established a technology of electric field measurement with an optical fiber sensor which uses an Electro-Optic crystal (EO probe). This technology enables electric field measurement in plasma source under beam acceleration without disturbing microwave electric field. In this study, first, validity of electric field measurement using the EO probe in the atmosphere was demonstrated by comparing experimental results with FDTD simulation. Then, we measured axial electric field distribution in the accelerated plasma. This experiments indicated that electric field distribution in the μ10 thruster was related to its beam current .