Hayabusa-2, Japan's second Asteroid Sample Return Mission
Hayabusa-2 is JAXA's (Japan Aerospace Exploration Agency) follow-on mission to the Hayabusa mission, the country's first round-trip asteroid mission that sent the Hayabusa (MUSES-C) spacecraft to retrieve samples of asteroid Itokawa. The initial Hayabusa mission launched in May 2003 and reached Itokawa in 2005; it returned samples of Itokawa — the first asteroid samples ever collected in space — in June 2010. Hayabusa means 'falcon' in Japanese. 1) 2) 3) 4)
The objective of the Hayabusa-2 sample return mission is to visit and explore the C-type asteroid 1999 JU3, a space body of about 920 m in length and of particular interest to researchers, because it consists of 4.5 billion-year-old material that has been altered very little. Measurements taken from Earth suggest that the asteroid's rock may have come into contact with water. The carbonaceous or C-type asteroid is expected to contain organic and hydrated minerals, making it different from Itokawa, which was a rocky S-type (stony composition) asteroid.
Table 1: Some background on the programmatic exploration of primitive bodies
Table 2: Objectives : Hayabusa vs Hayabusa-2 5)
Figure 1: Overall schedule of the Hayabusa-2 mission (image credit: JAXA)
Detailed information of asteroid 1999JU3 has been obtained by observations of ground-based telescopes. According to observation data of 2008, the diameter of 1999JU3 is estimated to be about 900 m, larger than that Itokawa, and the rotation period is around 7.6 hours. Observation of the reflected sunlight spectrum showed, that it has features of a C-type asteroid. It is rather difficult to determine the spin axis of asteroid 1999JU3 because of its rather spherical shape.
Figure 2 shows the orbit of asteroid 1999 JU3. The orbit is similar to that of Itokawa, and it is orbiting from just inside the orbit of the Earth to just outside the orbit of Mars. The inclination of the orbit is small like the one of Itokawa. Such an orbit is suitable for a small spacecraft like Hayabusa-2 to reach and return to Earth.
Legend to Figure 2: The blue circled lines in the figure illustrate the orbits of Mercury, Venus, Earth and Mars from inside, respectively, Itokawa's orbit is green, while the yellow orbit is that of 1999JU3.
The asteroid was discovered in 1999 by the LINEAR (Lincoln Near-Earth Asteroid Research) project, and given the provisional designation 1999JU3 (it hasn't been named so far). LINEAR is a collaboration of the United States Air Force, NASA, and MIT/LL (Massachusetts Institute of Technology /Lincoln Laboratory) for the systematic discovery and tracking of near-Earth asteroids.
Figure 3: Schematic of the science of Hayabusa and Hayabusa-2 missions (image credit: JAXA)
Project short history:
The Hayabusa-2 mission was proposed in 2006 at first. In this first proposal, the spacecraft was almost same as that of Hayabusa, because the project team wanted to start it as soon as possible. Of course, the team realized that parts had to be modified where trouble occurred in Hayabusa, but there were no major changes. The launch windows to go for launch to asteroid 1999 JU3 were in 2010 and 2011. However, JAXA could not start Hayabusa-2 mission immediately, because no budget was available. Hence, the launch opportunity was missed. The next launch window came up in 2014. Thus, the project postponed the launch date, and continued proposing Hayabusa-2. Since the launch was delayed, the project had time to change the spacecraft a little. New instruments were added, such as a Ka-band antenna and what is called "impactor." The project even calls Hayabusa-2 a new spacecraft. 6)
In May 2011, the status of Hayabusa2 project shifted to Phase-B, starting with the design of the spacecraft. In March 2012, the CDR (Critical Design Review) was done, and the team started manufacturing the flight model. The initial integration test started at the beginning of 2013, and the final integration test started at the end of 2013. At present (September 2014), the team has almost finished the final integration test, and the spacecraft will be shipped to the launch site soon.
International collaborations: The Hayabusa-2 mission involves international collaborations with Germany, the United States, and Australia. DLR (German Aerospace Center) and CNES (French Space Agency) are providing the small lander MASCOT. NASA was already a partner in the Hayabusa mission, a similar collaboration is under consideration for Hayabusa-2. The third collaboration is with Australia for capsule reentry as in the case of the Hayabusa mission.
Japan's Hayabusa-2 spacecraft is designed to study asteroid 1999 JU3 from multiple angles, using remote-sensing instruments, a lander and a rover. It will collect surface- and possibly also subsurface materials from the asteroid and return the samples to Earth in a capsule for analysis. The mission also aims to enhance the reliability of asteroid exploration technologies. 7) 8) 9)
In the current plan, the launch window for Hayabusa-2 is in late 2014. With this schedule, Hayabusa-2 will reach the asteroid in the middle of 2018, and return to the Earth at the end of 2020.
Figure 4: Artist's rendition of the Hayabusa-2 spacecraft (image credit: JAXA, Akihiro Ikeshita)
The Hayabusa-2 mission will utilize new technology while further confirming the deep space round-trip exploration technology by inheriting and improving the already verified knowhow established by Hayabusa to construct the basis for future deep-space exploration.
The configuration of Hayabusa-2 is basically the same as that of Hayabusa, with modifications of some parts by introducing novel technologies that evolved after the Hayabusa era. For example:
• The HGA (High Gain Antenna) for Hayabusa featured a parabolic shape, while Hayabusa-2 uses two planar HGAs with a considerably lower mass but with the same performance characteristics. The reason why Hayabusa-2 has two HGAs is that spacecraft has two communication links, Ka-band as well as the X-band links. In daily operations support, the team uses the X-band for data transmission, but for the download of the asteroid observation data, the Ka-band is used to take advantage of the higher data rate of 32 kbit/s, provided by the Ka-band link. The DDOR (Delta-Differential One-way Ranging) technique is used for very accurate plane-of-sky measurements of spacecraft position which complement existing line-of-sight ranging and Doppler measurements.
• The AOCS (Attitude and Orbit Control Subsystem) of Hayabusa-2 was improved, now featuring 4 reaction wheels for a more reliable service in case of need.
- During the cruise phase, Hayabusa-2 controls its attitude with only one reaction wheel to bias the momentum around the Z-axis of the body. This is to save the operating life of reaction wheels for other axes, because the project experienced that two reaction wheels of three equipped on Hayabusa were broken after the touchdown mission. 10)
- In this one wheel control mode, the angular momentum direction is slowly moved in the inertial space (generally called precession) due to the solar radiation torque. This attitude motion caused by the balance of the total angular momentum and solar radiation pressure is known to trace the Sun direction automatically with ellipsoidal and spiral motion around Sun direction. Based on this knowledge of the past, the attitude dynamics model for the Hayabusa-2 mission had been developed before the launch. According to the newly developed attitude dynamics model of Hayabusa-2, the precession trajectory is almost the ellipsoid around the attitude equilibrium point, and this equilibrium point is determined mainly by the phase angle around Z-axis of the body.
- In the actual operation of Hayabusa2, the spacecraft experienced already the one wheel control mode, and the attitude motion in this mode is nearly corresponding to the expected motion based on the dynamics model developed before the launch. The precession trajectory is ellipsoid around the equilibrium point, and the attitude dynamics model is verified by the actual flight data. In this one wheel operation, the Sun aspect angle is restricted within a certain limit angle in terms of the thermal condition of the spacecraft. Because the precession radius is determined by the initial attitude and the equilibrium point, the Sun aspect angle almost exceeds the limit angle due to the precession without change of the equilibrium point. At this operation, the project executes the attitude maneuver around the Z-axis to change the equilibrium point, in order to reduce the Sun aspect angle - and succeeded. After that, the project executed the maneuver again to change the equilibrium point to a close point in order to make the small precession trajectory (Ref. 10).
• IES (Ion Engine System) has been modified to account for the aging effect during extended support periods. The thrust level of IES was increased by 25%, using the same Xe microwave discharge ion engine system.
IES will be used for orbit maneuvers during the cruising of the Hayabusa-2's onward journey to the asteroid and return trip to Earth. The engine enables to make the round trip with only one tenth of the power consumption compared to that of chemical propellant.
Major improvements from the Hayabusa mission are:
- Countermeasures to plasma ignition malfunction of one ion source of an ion engine. Carefully coordinating each part of the ion engine to improve both ion source propulsion generation efficiency and ignition stability.
- Countermeasures to degradation and malfunction of three neutralizers that occurred after 10,000 to 15,000 hours of operation. To make the neutralizer's lifespan longer, the walls of the electric discharge chamber are protected from plasma and the magnetic field has been strengthened to decrease the voltage necessary for electron emission.
- The maximum power was successfully increased to 10 mN per ion engine from the conventional 8 mN.
Figure 5: Photo of the IES assembly (image credit: JAXA)
Figure 7: Bottom view of Hayabusa-2 spacecraft illustrating the various elements of the spacecraft (image credit: JAXA)
The Hayabusa-2 spacecraft has a stowed size of 1.6 m x 1 m x 1.25 m (height). With the solar panels deployed, the 600 kg satellite as a width of 6 m.
Table 3: Spacecraft system parameters
Figure 8: Photo of the Hayabusa-2 flight model, taken in Aug. 31, 2014, before shipping to the launch site, TNSC (Tanegashima Space Center). The SRC (Sample Return Capsule) is mounted to the bottom front side center of the spacecraft (image credit: JAXA)
Launch: The Hayabusa-2 spacecraft was launched on December 03, 2014 (04:22:04 UTC) on a H-IIA vehicle (No. 26) from TNSC (Tanegashima Space Center), Japan. The launch service provider was MHI (Mitsubishi Heavy Industries, Ltd). The launch was nominally and about 1 hr 47 minutes and 21 seconds after liftoff, the separation of the Hayabusa-2 spacecraft into an Earth-escape trajectory was confirmed. 11) 12)
• Shin'en-2, a nanosatellite technology demonstration mission (17 kg) of Kyushu Institute of Technology and Kagoshima University, Japan. The objective is to establish communication technologies with a long range as far as moon. Shin'en-2 carries into deep space an F1D digital store-and-forward transponder which offers an opportunity for earthbound radio amateurs to test the limits of their communication capabilities.
• ArtSat-2 (Art Satellite-2)/DESPATCH (Deep Space Amateur Troubadour's Challenge), a joint project of of Tama Art University and Tokyo University. DESPATCH is a microsatellite of ~30 kg. The microsatellite carries a "deep space sculpture" developed using a 3D printer, as well as an amateur radio payload and a CW beacon at 437.325 MHz.
• PROCYON (PRoximate Object Close flYby with Optical Navigation) is a microsatellite (67 kg) developed by the ISSL (Intelligent Space Systems Laboratory) of the University of Tokyo and JAXA. The objective is to demonstrate microsatellite bus technology for deep space exploration and proximity flyby to asteroids performing optical measurements. 13)
Orbit: The trajectory of Hayabusa2 for the whole mission is shown in the sun-earth fixed coordinate in Figure 9. The total cruising time is about 4.5 years, and the asteroid proximity period is about 1.5 years. So the total flight time is about 6 years. The departure C3 is 21 km2/s2, the total impulse of the ion engine is 2 km/s, and the reentry speed of the capsule is 11.6 km/s.
Legend to Figure 9: Hayabusa-2 is equipped with a high-specific impulse ion engine system to enable the round-trip mission. First one year after launch is an interplanetary cruise phase called EDVEGA (Electric Delta-V Earth Gravity Assist).
• On February 26, 2018, Hayabusa-2 saw its destination —asteroid Ryugu— for the first time! The photographs were captured by the ONC-T (Optical Navigation Camera - Telescopic) imager onboard the spacecraft (Figure 10). Images were taken between noon JST on February 26th and 9:00am the following morning, with about 300 shots taken in total. Data for nine of these images were transmitted from the spacecraft on February 27th, allowing us to confirm that Ryugu had indeed been seen. The animation shows these nine consecutive frames. 14) 15)
- Ryugu's brightness from Hayabusa-2 is about magnitude 9, which would be impossible to see with the naked eye but visible with the ONC-T. Looking at the image above, you can see how the position of the surrounding stars relative to Ryugu appears to change as Hayabusa-2 moves towards the asteroid. The distance between Ryugu and Hayabusa-2 when the images were taken is about 1.3 million km. Ryugu as seen from Hayabusa-2 is in the direction of the constellation Pisces.
- Ryugu was photographed when the Sun, Hayabusa-2 and Ryugu were almost in a line. This configuration can be seen in Figure 11, which shows a snapshot of the header from the Hayabusa-2 website on February 26, 2018, which continuously updates to show the position of Hayabusa-2. If you were to stand on Ryugu, Hayabusa-2 would be seen in the direction of the Sun.
- Hayabusa-2 is currently using its ion engine to make adjustments to its course. This makes it difficult to alter the orientation of the spacecraft. However, at the alignment shown in Figure 11, the ONC-T camera can image Ryugu without needing to make significant changes to the spacecraft's orientation. This made February 26th the perfect time to try and capture Ryugu's image with the ONC-T. From the data, Ryugu was observed to be exactly at the expected location based on Hayabusa-2's estimated position. This tells us that Hayabusa-2 is flying on the planned course.
- "Now that we see Ryugu, the Hayabusa-2 project has shifted to the final preparation stage for arrival at the asteroid. There are no problems with the route towards Ryugu or the performance of the spacecraft, and we will be proceeding with maximum thrust," explains Project Manager, Yuichi Tsuda.
- The remaining images will be transmitted back to Earth from the spacecraft and allow us to further confirm the asteroid and spacecraft location. Although we can currently see Ryugu only as a point, it is very exciting for the whole project team to catch sight of the destination!
• As of September 2017, two of three IES (Ion Engine System) sessions have been completed. The first and second IES session, 798 hours and 2558 hours of IES operation was executed, respectively. The total IES operation time is about 3,900 hours and a velocity increment of about 580 m/s has been obtained. 16)
- In December 2017, the last IES session will be started and will continue until June 2018. After the completion of the third IES operation, the Hayabusa-2 spacecraft will arrive in the vicinity of Ryugu. Since there are large uncertainties on the asteroid ephemeris, the Hayabusa-2 spacecraft will perform an optical navigation campaign to reduce the relative position error to the asteroid. The direction of the asteroid will be determined by the ONC-T (Optical Navigation Camera - Telescopic) imager; if required, a series of guidance maneuvers with the RCS will be used to correct the trajectory toward the asteroid. A month of an optical navigation campaign is scheduled for this final approach phase.
- Hayabusa-2 will arrive at the asteroid vicinity in the summer of 2018. Hayabusa-2 will explore Ryugu for 1.5 years and return to the Earth in the winter of 2020. The 1.5 year allocation for proximity phase operations is much larger than the 3 months of Hayabusa's proximity operation, however the schedule is marginal. Three touchdown operations need to be planned within the limited amount of time to satisfy the mission success criteria.
- There are three main uncertainty parameters, which impact the schedule. Those are: gravity, thermal environment, and spin axis orientation of Ryugu.
- Firstly, the Ryugu gravity limits the number of descent operations. This is because the limitation of the fuel onboard and the fuel consumption per descent operation. The smaller the gravity constant, the less fuel requirement - thus, more descent attempts can be done. Currently, the porject assumes the gravity coefficient between 11 to 92 m3/s2. The "worst case" for the scheduling is the largest number of the descent operations, derived by the smallest gravity. The maximum number of descent operations is considered as a slot in the nominal schedule.
- Secondly, the thermal environment of the asteroid impacts the schedule. This constraint is the solar distance constraint for the TD (Touch Down) operation. There are upper constraints in the solar distance because the spacecraft temperature becomes too high when the spacecraft approaches the asteroid surface and receives the thermal inputs both from the asteroid as well as from the Sun. The current best estimate of Ryugu's thermal environment derived resulted in a touch down epoch requirement between October 2018 and May 2019; this corresponds with a 1.25 AU sun distance. This constraint will be updated after the Ryugu in-situ observation of Hayabusa-2 in the Summer of 2018.
- The third uncertainty parameter is the spin axis orientation of the asteroid. Although there is series of light curve observation of Ryugu from ground based telescope, the spin axis orientation is still uncertain. The spin axis impacts the Hayabusa-2 visibility of the asteroid. Since the Hayabusa-2 spacecraft stays in the vicinity of the Earth-asteroid line, the limited visible latitude is given as a function of spin axis. Assuming that the asteroid is a single spinner, the visible latitude varies due to the asteroid orbital motion around the Sun. The orbital motion changes the geometry between the asteroid and the Earth, therefore the visible latitude changes. — The accessible region for the descent operation also depends on the spin axis orientation. Due to the GNC limitation, the descent operation need to be done within 200 m from the asteroid-Earth line. This restriction describes the accessible latitude band as a function of time. Figure 12 illustrates the schematic image of the change of the visible area and accessible latitude band. The relative position in HP (Home Position) frame will change due to the orbital motion around the Sun.
- Figure 13 describes the latitude band and visibility of a sample spin axis orientation. The blue region is the touch down or MASCOT release point accessible latitude band. The gray zone during the end of 2018 is the conjunction phase and neither observation nor critical operation can be performed. In this example, the accessible latitude before and after the conjunction have a large difference. 17)
- Considering all three points, the nominal schedule is conducted as shown in Figure 14. The TD1 is allocated just before the conjunction (with margin). All the previous operations are derived by the TD1 operation. Especially, the landing site selection is the most important input for the operation.
Legend to Figure 14: LSS =Landing Site Selection, SCI =Small Carry-on Impactor.
• As of September 2016, 1346 hours of the ion engine operation have been achieved as planned. Three touch downs/sample collections, one kinetic impact/crater generation, four surface rovers deployment and many other in-situ observations are planned in the asteroid proximity phase. The operation team will perform extensive operation practice/rehearsal using a hardware-in-the-loop simulator in the year 2017 to be ready for the asteroid arrival in the summer 2018. 18)
- Hayabusa-2 adopts a novel attitude control technique called "Solar Sail Mode" for reliable and fuel-efficient attitude operation. 19) It is used in ballistic cruise periods and realizes a coarse Sun pointing attitude using one RW only and actively utilizing SRP (Solar Radiation Pressure). The last 1 month of the commissioning phase from January 24 until March 2 was dedicated to the SRP torque characterization and Solar Sail Mode attitude control testing/demonstration, which was completed successfully.
- Forward Cruise Operation: Hayabusa2 started the forward cruise phase (cruise from Earth to Ryugu) on March 3, 2015, following the completion of the commissioning phase. Total of approximately 7000 hours IES operation is planned to reach Ryugu. Table 4 shows the history of the IES operation from launch until present. Figure 15 shows the trajectory from EGA to Ryugu arrival with the IES thrust vector history.
- Ballistic flight phase (cruise without IES operation) is effectively utilized for precise navigation, regular onboard instruments checkout/calibrations, and engineering demonstration/testing. For example, Mars observation campaign was conducted from May 28-June 9, 2016 at the closest distance (41 million km) for ONC/TIR/NIRS3 calibration purpose. The Ka-band operation between Hayabusa2 and an ESA's deep space station (ESTRACK Malargue station in Argentina) was conducted successfully on July 5-8, 2016. The Ka-band DDOR was tested several times using NASA(DSN) and ESA(ESTRACK) stations and achieved the world-first successful inter-agencies operation of the Ka-band DDOR under the CCSDS (Consultative Committee for Space Data Systems) DDOR operation standard. Besides testing, X-band DDOR is regularly used for precise orbit determination in ballistic cruise periods.
- EGA (Earth Gravity Assist) operation: Hayabusa2 performed the EGA on December 3, 2015. Four TCMs (Trajectory Correction Maneuvers) were planned and three were actually conducted. The final TCM was canceled because the spacecraft had already been guided accurately with the three TCMs. Table 4 shows the list of TCMs for the gravity assist with the guidance accuracy results.
- The Earth closest approach occurred at 10:08:07 UT, December 3, 2015 over the Pacific Ocean (Figure 16). The closest distance was 3090 km (from the Earth surface) and the flyby deflection angle (angle between incoming and outgoing velocity vector in ECI frame) is 83º. The interplanetary velocity increment by this EGA is 1.6 km/s. The spacecraft experienced an eclipse for 20 minutes, which is the longest battery-powered operation for Hayabusa-2 throughout its mission.
- The operation team is now preparing for operation practice/rehearsals using the hardware-in-the-loop simulator in the year 2017 to be ready for the asteroid arrival in the summer 2018.
• July 14, 2016: MASCOT (Mobile Asteroid Surface Scout) of DLR has been travelling on board the Japanese Hayabusa2 spacecraft for the last 1.5 years, and is currently at approximately 65 million km from Earth. On 14 July 2016, DLR (German Aerospace Center) engineers in the LCC (Lander Control Center) in Cologne switched the shoebox-sized lander and its four German and French-built instruments back on, and will spend the next few days finding answers to two questions: How is MASCOT's state of health? And how are the experiments on board? "We do a check-up once a year to find out whether all system components and instruments are still in good working order," explains Christian Krause from the Lander Control Center team at DLR. 20)
- The Hayabusa2 spacecraft by JAXA (Japan Aerospace Exploration Agency) set off on its mission on 3 December 2014, carrying the French-German MASCOT lander. One year later, the duo zipped round the Earth to gain momentum and sent back photos from our planet before continuing on toward the asteroid Ryugu. The spacecraft will venture deeper into space until the summer of 2018, when it will enter orbit around the celestial body that DLR planetary researcher Ralf Jaumann refers to as a 'beautifully primitive object'. "During this mission, we will be investigating primordial material from the solar nebula; it has remained practically unchanged in its 4.5 billion years of existence." Then, while the Hayabusa2 spacecraft measures and analyses the asteroid from its position in orbit, MASCOT will descend to its surface to conduct scientific measurements. The Japanese spacecraft will also take on soil samples that it will bring back to Earth in 2020. "This is a complete package. There has never been anything like this before: we will be observing and mapping remotely, measuring the asteroid, analyzing its surface and bringing the samples back to Earth." But this complete package requires the concerted efforts of engineers and scientists from Germany, France and Japan who have joined forces in an international cooperation.
• December 25, 2015: Before and after the Earth swing-by, the laser altimeter (LIDAR) on Hayabusa2 attempted to receive laser light from the SLR (Satellite Laser Ranging) ground stations. 21)
- After the swing-by, the Mt. Stromlo station at SERC (Space Environment Research Center Australia) in the suburbs of Canberra, Australia, transmitted laser light towards Hayabusa-2. The spacecraft successfully received the beam using the onboard LIDAR that can send and receive laser signals to accurately establish the range of objects from the spacecraft. At the time of the transmission from Mt. Stomlo, Hayabusa-2 was 6,700,000 km from Earth. This success established the one-way 'up link' of the optical connection.
• December 14, 2015: JAXA confirmed that the Asteroid Explorer "Hayabusa-2" is cruising on its target orbit after measuring and calculating the post-Earth-swing-by orbit. With the swing-by results, the explorer's orbit turned by about 80 degrees and its speed increased by about 1.6 km/s to about 31.9 km/s (against the sun), thus the orbit achieved the required target velocity (Figure 17). According to the operational support provided by the NASA DSN (Deep Space Network) stations and the ESA (European Space Agency) deep space ground station Malargüe in Australia, the Hayabusa2 is in good health. 22) 23)
- At 9:00 UTC on Dec. 14, 2015, the Hayabusa-2 is at a distance of ~4.15 million km from Earth, and about 144.85 million km from the sun. Its cruising speed is 32.31 km/s (against the sun). Hayabusa-2 is increasing its speed under the influence of the sun's gravity after the swing-by.
- After the swing-by, Hayabusa-2 took images of the Earth using its onboard ONC-T (Optical Navigation Camera - Telescopic). The ONC-T can shoot color images using seven filters. The image of Figure 18 is composed by using three of these filters. One can see the Australian continent and Antarctica in the image. The South Pole is not lit by the sun during the(northern hemisphere) summer, and the meteorological satellites also do not cover the Antarctic continent to take their imagery, hence the shot this time is precious.
Figure 18: Shot at 22:09 (UTC) on Dec. 4, 2015, about 340,000 km from the center of the Earth One can see the Australian continent on the upper right, and Antarctica on the lower right (image credit: JAXA)
• December 3, 2015: JAXA performed an Earth swing-by operation of the Asteroid Explorer "Hayabusa-2 " in the evening of December 3 (Thursday), 2015 JST (Japan Standard Time). The Hayabusa-2 flew closest to the Earth at 10:08 (UTC) and passed over Hawaii at an altitude of about 3,090 km. - The Hayabusa-2 project team is currently measuring and calculating the post-swing-by orbit. It will take about a week to confirm if the explorer entered the target orbit. We will report the result once it is determined. The Hayabusa-2 is in good health. 24)
- ESA's deep-space Malargüe ground station in Australia lent a helping ear as Japan's Hayabusa-2 asteroid mission visited Earth on Thursday. 25)
• October 5, 2015: The asteroid 1999 JU3, a target of the Asteroid Explorer "Hayabusa-2," was named "Ryugu". 26)
- JAXA conducted a naming proposal campaign between July 22 and August 31, 2015.
- Selection reasons: In the Japanese ancient story "Urashima Taro", the main character, Taro Urashima, brought back a casket from the Dragon's palace, or the "Ryugu" Castle, at the bottom of ocean. The Hayabusa2 will also bring back a capsule with samples, thus the theme of "bringing back a treasure" is common.
- Rocks containing water are expected to exist on asteroid 1999 JU3. The name "Ryugu" also reminds us of water, as "Ryugu Castle" is under the ocean. The name is not similar or identical with any other already existing names of planets or asteroids, and there were many entries for this name among suggested names that are related to mythology.
- According to the naming rule stipulated by the International Asteroid Union (IAU), the name "is preferably from mythology" and the "Ryugu" fits that rule. Also, there is little concern of infringing the Trademark Law or any other third party trademarks.
• June 8, 2015: The Hayabusa-2 spacecraft has been continuously operating its ion engine for the second time since June 2, and successfully completed its operations at 0:25 a.m. on June 7 (Japan Standard Time.) The second continuous operation lasted for 102 hours as scheduled. 27)
- The ion engine boost was needed in preparation for the Earth swing-by planned in December 2015, and the total hours of the first and second operations (409 hours and 102 hours respectively) reached 511 hours.
• April 29, 2015: ESA is set to support Japan's ‘touch-and-go' Hayabusa-2 spacecraft, now en route to a little-known asteroid, helping to boost the scientific return from this audacious mission. A flawless launch last December marked the start of a six-year round-trip for Japan's Hayabusa-2, which is on course to arrive at the carbon-rich asteroid 1999 JU3 in June 2018. 28)
- In the first such support provided to a Japanese deep-space mission, ESA's 35 m diameter dish at Malargüe, Argentina, will provide up to 400 hours of tracking, establishing radio contact as the asteroid arcs through the Solar System between 135 million to 210 million km from the Sun. Telecommands from mission controllers at JAXA (Japan Aerospace and Exploration Agency) will be fed to the station via ESA/ESOC in Darmstadt, Germany.
- ESA is using its ESTRACK network with deep space stations at Malargüe (Argentina), Cebreros (Spain) and New Norcia (Australia). On 22 April, ESA completed a live, inflight compatibility test, linking the Malargüe station with the Japanese spacecraft, demonstrating the network's readiness to provide tracking for the Hayabusa-2 mission.
• April 10, 2015: Hayabusa-2 is stably flying in space. The new fiscal year has just started in Japan, and JAXA is taking a new step as we became a 'National Research and Development Agency' from the previous independent administrative agency. The Hayabusa-2 project is also taking a fresh step with a new team, including handing the baton over to a new project manager, namely from Hitoshi Kuninaka to Yuichi Tsuda. All members of the project are engaged in the mission with a fresh mindset. 29)
• March 18, 2015: LIDAR (LIght Detection And Ranging) is one of the satellite buses onboard Hayabusa-2. Basically LIDAR is used to measure the distance between the satellite and the target asteroid, 1999JU3. The distance is important information to control the satellite in descending operation such as close-up observation from low altitude and separation of rover, lander, and small carry on impactor. Especially an accurate measurement of the distance is required during touch-down and sampling operations. Thus, the LIDAR onboard Hayabusa-2 is required to measure about 4 orders of magnitude variety of distance, from 30m to 25km. To satisfy this requirement, the LIDAR has two types of receiver optics, one is for near range (nearer than 1km) and another is for far range (farther than 1km). 30) 31)
- The distance from the target asteroid, 1999JU3, is one of the fundamental information for operational and scientific observations. For example, the gravitational acceleration is calculated from the change rate of the distance from the target asteroid during free fall operation. Construction of a shape model requires a characteristic length scale in camera image. And also, the measurement of the absolute reflectance from camera image and/or spectrometer data requires an absolute distance from the target.
- The dust count mode is one of the operational modes of the LIDAR in which the LIDAR detects faint scattered light from dust grains on the line of sight. The distribution of dust grains along the line of sight is estimated from the profile of received light and the effective reflectance of dust cloud. Its function was tested both electrically and optically, and it seemed working properly.
Figure 19: Schematic view of dust count mode. Thick curve in the bottom figure represents hypothetical profile of received light. Black-and-white line at the bottom represent an output of dust count mode observation (white=0, black=1), image credit: Hayabusa-2 team
• March 5, 2015: Hayabusa-2, launched on Dec. 3, 2014, completed its initial functional confirmation period of about three months. The explorer was moving to the cruising phase on March 3 while heading to the asteroid 1999 JU3. 32)
- The Hayabusa2 is in good health. It will be under preparatory operation including speed increase by continuous operation of the ion engines for an Earth swing-by scheduled in Nov. or Dec., 2015.
- The project plans to increase the cruising speed of the explorer (60 m/s) by operating two ion engines twice (in total about 600 hours or 25 days) until the Earth swing-by. For the first operation, the project will gradually increase the time duration of continuous ion-engine operation from March 3 onwards, and will operate the engines for about 400 hours within March. The second operation is scheduled in early June.
• Feb. 3, 2015: Hayabusa-2 is now undergoing the initial functional confirmation. Basic operations and performance of onboard instruments and ground systems have been tested one by one as of the end of January. Some examples: 33) 34)
- Ion engine test operation (one unit at a time). Four ion engines were being operated one by one. A thrust of 7-10 mN was generated on the orbit for the first time.
- Establishing communication by Ka-band communication equipment (between Jan. 5 to 10, 2015). Communication was successful between the Hayabusa-2 and NASA DSN stations to establish deep-space Ka-band communication for the first time for a Japanese space explorer. Ka-band communication will be used to send observation data during the mission for the Hayabusa-2 to stay near the asteroid.
- Ion engine can autonomously operate for 24 hours. Long duration of autonomous operation with two or three ion engines was tested, and 24-hour continuous operation was attained. The maximum thrust was confirmed to be about 28 mN, which is the expected value.
• Dec. 5, 2014: JAXA) confirmed the completion of a sequence of the important operations for the Asteroid Exploration Hayabusa-2 mission including the deployment of the horn part of the sampler that captures samples from the asteroid's surface, the release of the locks for the launch that ratchet the gimbal that controls the direction of the ion engine, and functional verification of the three-axis stabilization controls and the ground precision orbit determination system. With this confirmation, the critical operation phase of the Hayabusa2 was completed. 35)
• JAXA received the first signals of Hayabusa-2 on December 3, 2014 at 6.44 UTC on Dec. 3, 2014, acquired by the NASA DSN (Deep Space Network) at Goldstone in California; it was confirmed that its initial sequence of operations including the solar array paddle deployment and sun acquisition control have been performed normally. - The explorer is now in a stable condition. 36)
Sensor complement: (SMP, SCI, NIRS3, TIR, MINERVA-II-1A/1B/2, DCAM3, ONC-T, MASCOT, SRC)
SMP (Sampler Horn):
The SMP, also referred to as "sampler mechanism", will collect samples from the surface of the asteroid. The basic design is the same of that aboard the Hayabusa mission; thus, the same mechanism will be used, which is a small projectile to be shot as soon as the tip of a cylinder-shaped horn touches the asteroid surface, then materials ejected from the surface will be collected in a catcher (storage).
Major improvements from the Hayabusa:
• The seal performance has been improved for Hayabusa-2, and a newly developed metal seal method is applied, so that volatile gas such as rare gas can be brought back thanks to high airtightness.
• The number of compartments for the catcher that will store captured materials is increased by one to three compartments compared to two for the Hayabusa.
• For the Hayabusa2, the edge of the sampler horn was folded back inward as you can see in the left figure so that sand gravel can be hung on the cuff (gravel of 1 to 5 mm can be captured.) If the explorer makes an emergency stop during its ascending, gravel will keep going up to be stored in the catcher. This mechanism is a backup for sample retaining. 37)
Figure 20: Photo of the SMP assembly (image credit: JAXA)
SCI (Small Carry-on Impactor):
SCI is a new payload, an impactor caller "Liner", which is featured on Hayabusa-2. It consists of a small box, about 30 cm in size, which will be released a few hundred meters above the surface of the asteroid. The 2 kg copper liner will be shot at high speed (2 km/s) into the surface of the asteroid to create a small artificial crater by collision. The size of the crater may be about 2 or 3 m in diameter. The purpose of this impact experiment is not only to study the physical characteristics of the asteroid surface, but also to reveal the subsurface material, of which Hayabusa-2 will try to recover samples. It is expected that the fresh crater samples are less weathered by the space environment or heat. — The impactor (Liner) is made of pure copper to be able to easily identify the sample from other materials on the asteroid.
Figure 21: SCI pyrotechnics: Left: A conical shape structure filled with explosives. The "Liner" will be ejected forward at a high speed by explosive power. Right: The flying "Liner" at 2 km/s (image credit: JAXA)
NIRS3 (Near InfraRed Spectrometer, and 3 comes from "3µm") and TIR (Thermal Infrared Imager)
The Hayabusa2 will be orbiting asteroid 1999JU3 at an altitude of ~20 km, and observe the asteroid by remote sensing. Two kinds of infrared observations will play an important role in this remote sensing.
• One is the near infrared spectrometer "NIRS3" to investigate mineral and water metamorphism or reciprocal chemical action of minerals and water by spectroscopic observations with near infrared rays.
• The other is the TIR (Thermal Infrared Imager) to study the temperature and thermal inertia of the asteroid, by capturing images of thermal radiation from the asteroid. In other words, it will find particles of the soil and the porosity of the mass of rock that influence the temperature.
Figure 22: Photo of the NIRS3 (left) and of the TIR (right) instruments (image credit: JAXA)
MINERVA-II (Rover payload)
There are three small MINERVA-II rovers in total. They are of the heritage of MINERVA of Hayabusa, and they move on the surface of the asteroid by hopping. Hayabusa had only one MINERVA rover which failed to land it on the surface of Itokawa. Hayabusa-2 has two kinds of rovers, MINERVA-II-1 which separates into two (A and B), and MINERVA-II-2, so there are three small rovers in total. The rovers have small sensors (cameras, thermometers) onboard to convey their information to the orbiting Hayabusa-2 spacecraft.
Figure 23: Schematic view of the MINERVA-II rovers (image credit: JAXA)
All the rovers have hopping mobile systems to explore over the surface after the deployment from the mother spacecraft at the vicinity of the target asteroid. 38)
Table 5: Specification of the twin rovers (Rover-1A and Rover-1B)
DCAM3 (Deployable Camera)
DCAM3 is a miniaturized optical camera with the objective to observe the artificial impact of SCI ( Small Carry-on Impactor) on the asteroid. DCAM3 takes images of the spreading ejecta motion on the asteroid, which provides information of surface physical properties and ejecta behavior under microgravity. 39)
Collisions between primitive planetary bodies are one of the most important physical processes in the planetary accretion from planetesimals to planets. DCAM3 is a detachable camera inherited from DCAM-1,2 of the Japanese IKAROS mission. A separable instrument is necessary to obtain close up views of the impact, because the mother ship (Hayabusa-2) will be hiding in a safe region far from the impact point to avoid a risk that the mother ship encounters high-speed ejecta from the asteroid during the impact operation.
The impact observation part in the DCAM3 system consist of a separable small camera cylinder, its detachment instrument, and receiving antennas on the mother ship (Figure 24). After separation in the SCI impact operation, DCAM3 starts its observation at visible wavelength and sends image data on one-way communication to the mother ship. The design of DCAM3 is similar to IKAROS DCAM1,2 (Figure 25), but has an additional high-resolution camera in the body.
The DCAM3 instrumentation is provided by a Japanese University Consortium in cooperation with JAXA. The University Consortium consists of: University of Tokyo, Kashiwa; Kobe University, Kobe; JAXA (Japan Aerospace Exploration Agency), Sagamihara; Planetary Exploration Research Center, Chiba Institute of Technology, Narashino; Kochi University, Kochi; The Graduate University for Advanced Studies, Sagamihara; and the University of Occupational and Environmental Health, Kitakyusyu.
The scientific observations are performed by DCAM3-D, a wide-angle high-resolution camera and its fast digital transmission component in the DCAM3 body. The scientific objectives of this camera are summarized as (1) clarifying the subsurface structure, and (2) constructing the impact scaling rule applicable to the surface of asteroid 1999JU3. The observation objects of the camera are ejecta and a subsequent crater of the SCI impact, and a relative position of the SCI to the asteroid when it is launched. DCAM3-D can determine the size and the angle of the ejecta curtain, and the speed of the ejecta spreading or fragment spattering, which are the key information for the above objectives. In addition, low-speed ejecta (dust) spreading will possibly be observed around the DCAM3 in a few hours after the impact.
The DCAM3-D CMOS detector produces 2000 x 2000 pixels 8 bit monochromatic images with a 74º x 74º wide-angle optics. It takes 1 frame/s sequential images at maximum. Figure 26 shows a virtual image of spreading ejecta on asteroid 1999JU3 taken by the DCAM3-D in an ideal position. The optical camera has enough-high space resolution and bright resolution for its sciences. DCAM3 continues to produce data for a few hours until the batteries run out or the DCAM3 falls and crashes on the asteroid. The digital communication device can send the image data to the mother ship with 4 Mbit/s at maximum. Instruments in the mother ship store all data taken by the deployed camera. The total volume of image data is estimated to be approximately 5 Gbit after compression.
Figure 26: A virtual image of asteroid 1999JU3 and ejecta cone taken by the DCAM3-D camera. The blue oval sphere shows a potential region in which SCI exists. The green cone shows a possible trajectory of the projectile (image credit: University Consortium, JAXA)
ONC-T (Optical Navigation Camera-Telescopic) imager
The ONC-T was developed under collaboration between JAXA, the University of Tokyo, Kochi University, Rikkyo University, Nagoya University, Chiba Institute of Technology, Meiji University, The University of Aizu, the National Institute of Advanced Industrial Science and Technology (AIST).
The ONC system onboard the Hayabusa-2 spacecraft consists of one telescopic camera (T) and two wide-angle cameras (W1 and W2). ONC-T is a telescopic camera with seven band-pass filters in the visible and near-infrared range. These filters are placed on a wheel, which rotates to put a selected filter for different observations, enabling multiband imaging. 40)
The main objective of this instrument is to optically navigate the spacecraft to asteroid Ryugu (1999 JU3) and to conduct multi-band mapping the asteroid for choosing touchdown candidate sites and understanding the nature of this asteroid.
Figure 27: Photo of the ONC-T imager (image credit: JAXA)
MASCOT (Mobile Asteroid surface SCOut)
MASCOT is a small lander provided by DLR (German Aerospace Center) in collaboration with CNES (French Space Agency). Building upon the successful joint development of the Philae lander of the Rosetta mission, DLR and CNES have studied this lander since 2008, and its realization was formally decided in early 2012. 41) 42) 43) 44)
After Hayabusa 2's arrival to its target – asteroid 1999JU3 - in 2018, MASCOT will be dropped to the surface where it will perform in-situ investigations for about 15 hours. A hopping mechanism of MASCOT will enable measurements at several locations. The payload of MASCOT consists of four scientific instruments, a wide angle camera, an imaging IR spectrometer (MicrOmega), a radiometer (MARA) and a magnetometer. MASCOT will significantly enhance the overall scientific return of the Hayabusa-2 mission by providing context measurements to support the interpretation of analyses of the returned samples. 45)
The general concept of MASCOT is to provide a small landing system intended to be deployed from a main spacecraft (or "mother-ship") on an asteroid sample return mission. MASCOT has been specifically designed to be compatible with JAXA Hayabusa-2 spacecraft and the environment given by the target asteroid 162173 Ryugu to fulfill the following scientific objectives: 46)
- In-situ observations of undisturbed materials and microscopic scale observations, not possible from the main spacecraft, to determine the asteroid regolith structure, texture and composition
- Magnetic properties measured during the descent, during the hopping maneuvers and on the surface on more than one site
- Temperatures of one asteroid rotation and thermal properties on more than one site
- Help the selection of the sampling spot(s) by the main spacecraft.
MASCOT is a 11 kg lander with a 3 kg payload, of 0.3 x 0.3 x 0.2 m3 volume and a single energy source (of 220 Wh) allowing about a 12 hours of mission duration. The design for MASCOT is based on the following subsystems:
Figure 28: Photo of the MASCOT flight model (image credit: DLR, Telespazio VEGA)
• Structure & Accommodation: The lander baseline design features a highly integrated carbon-fiber composite structure with a middle wall as main load bearing element. A common electronics box, called E-box, houses all the electronics. The lander is connected with the main spacecraft Hayabusa-2 via MESS (Mechanical and Electronics Support Structure), which comprises the PRM (Preload Release Mechanism) and the separation device. An umbilical connector supplies the power during the cruise phase.
• Thermal Control: The thermal control concept is based mainly on passive means such as MLI (Multi-Layer Insulation) and color coatings. Thus, it relies also on a day-night-cycle on the asteroid's surface to balance the heat-load build-up. This constraint drives the selection of the operating asteroid site for MASCOT. An exception to the passive concept are the heaters used for the thermal control of batteries and MicrOmega during the cruise and commissioning phase. The heater is controlled by Hayabusa-2.
Figure 29: Overview of MASCOT elements (image credit: DLR, Telespazio VEGA)
• Power: The power and energy supply is maintained by a primary battery, provided by the SAFT company (France), based on an earlier development for Philae. The unregulated bus voltage is converted into auxiliary voltages for MASCOT electronics by an internal PCDU (Power Control and Distribution Unit) located in the E-box. During the cruise and commissioning phase, the power will be supplied by the Hayabusa-2 spacecraft via a regulated power line. The power subsystem is a French contribution to MASCOT.
- During the cruise phase, the power is supplied by the Hayabusa-2 spacecraft via a dedicated power line, through the MESS connector. After separation, the power is supplied by a primary battery. Its capacity is dimensioned for an operating time of two asteroid days, ~16 hours nominal. Both the Hayabusa-2 power and the battery power line is converted into auxiliary voltages for MASCOT electronics by an internal PCDU (Power Control and Distribution Unit), located in the common E-box. The PCDU is made of two similar units forming a redundant architecture, with redundant power converters for the secondary voltages to redundant and non-redundant equipment. The data link with the OBC is via two UART RS422 lines for each PCDU unit.
• GNC (Guidance and Navigation) sensors: The GNC sensors are used to determine the MASCOT motion state and orientation on the asteroid's surface. They are distributed all over the MASCOT body and are acquired via the hot redundant A/D converter of the OBC. The algorithms to determine the motion status, the illumination status and the MASCOT side in contact with the soil (side2soil) are implemented in the MASCOT OBSW. The navigation sensors consist of:
- OPS (Optical Position Sensors): 5 optical distance sensors are mounted on the outer skin of MASCOT on each side, except -Z, the nominal side at surface contact. The optical position sensors use an IR LED and a photodiode. The IR LED is switched on/off by the OBC to perform a differential measurement and detect the presence of a surface. With these sensors it is also possible to detect if MASCOT is moving or is lying still, the side on which MASCOT is lying and the distance to the soil.
- PECs (Photo-Electric Cells): These sensors are used for redundancy to the OPS. One PEC is mounted on each side of MASCOT. There are a total of 6 sensors. The output of the PEC is a voltage proportional to the cosine of angle between the sun vector and the cells normal vector.
- Separation sensor: Break-wire contacts on the power connector from Hayabusa-2, integrated in the MESS, will detect the separation.
• OBC (On-Board Computer): The MASCOT OBC is a dual redundant computer composed by 2 CPU boards (CPU-M, CPU-R) to execute the OBSW (On-Board Software) and 2 I/O boards (IOM-M, IOM-R) to provide interfaces with all MASCOT subsystems. The 4 PCBs are accommodated in the common E-box. The two CPU boards work in cold redundancy, while the I/O boards work in warm redundancy, cross-strapped with the CPU boards via SpaceWire links. Each I/O board uses a radiation tolerant FPGA to implement:
- the I/O controllers for the external UART serial interfaces
- the I/O controllers to handle the analog section and bi-level I/O interfaces
- a NAND Flash controller featuring a page level Reed - Solomon encoder / decoder
- the CPU switch-over logic driving the reconfiguration of the CPU from main to redundant or vice versa
- a local timer to centrally maintain the MOBT (MASCOT On-Board Time) independent from the CPU
- a small RAM area, EDAC protected, used as Safe Guard Memory (SGM) to propagate the OBSW context in case of CPU reset or switch-over.
These devices are accessible by the two CPU board via SpaceWire interface. The I/O board also contains the NAND Flash mass memory device, the analog section circuits, the RS422 transceivers, the SpaceWire LVDS transceivers and other interface drivers/receivers needed for the external interfaces. The CPU board uses an Aeroflex GR712RC SOC (System On a Chip) clocked at 40 MHz, embedding a dual LEON3FT CPU core and a few other cores, like SpaceWire controllers, timers, DSU (Debug Support Unit) etc. The board also accommodates the program and data RAM (implemented as a SRAM multi chip module), a reprogrammable non-volatile memory PROM (implemented as a MRAM device), the Watch Dog and Reset circuits and the SpaceWire LVDS and other transceivers needed for the external interfaces. The SOC DSU JTAG interface is routed to the E-box test connector through dedicated transceivers. The SpaceWire serial interfaces to the MicrOmega and CAM instruments are implemented by the CPU board.
In terms of processing and memory resources, each OBC section (1 CPU + 1 IOM) provides:
- CPU speed: minimum 40 DMIPS, 10 MFLOPS
- RAM (SRAM): 16 MByte EDAC protected
- PROM (MRAM): 256 kByte / 204 kByte with EDAC enabled
- NAND Flash: 1 GByte / 860 MByte with Reed-Solomon encoding.
• RF communications: The communication architecture is based on a redundant CCOM (Child -COM) transceiver provided by JAXA; it is identical with the communications subsystem of the MINERVA small landers. The data link is based on CCSDS TM/TC packets and the transmission protocol between the MASCOT CCOM and the Hayabusa-2 PCOM (Parent-COM) ensures one-to-one delivery of the CCSDS packets by automatic retransmission requests in both directions in case of errors. MASCOT is equipped with two CCOM units working in cold redundancy. Each CCOM unit has two RF ports connected to two patch antennas via an RF coupler. One antenna located on top (+Z) and the other on the bottom (-Z) of MASCOT ensure almost omnidirectional coverage.
The maximum bit rate possible is ~37.04 kbit/s for telemetry packet downlink; ~1.7 kbit/s for telecommand packet uplink. An UART RS422 TM/TC umbilical interface is foreseen through the E-box test connector, to be used as alternative to the CCOM RF channel only during ground AIV activities. After the Hayabusa-2 launch, only the CCOM RF channel is used for TM/TC with MASCOT, via the Hayabusa-2 data handling unit (DHU) and the lander's computer,OME (On-board Minerva Equipment).
MASCOT antenna: For cruise phase operations, a dedicated MESS antenna is required whereas for on asteroid operations, the OME-A Hayabusa antenna is used. Two omnidirectional antennas are positioned on the top and bottom sides of MASCOT (Figure 30). The communication subsystem (CCOM transceiver excepted) is a French contribution to MASCOT.
Inside MASCOT a redundant set of JAXA-provided Child-Communication-transceivers (CCOM) is used to communicate with its counterpart - the Parent-Communication transceiver (PCOM) - on board of Hayabusa2 based on a half-duplex communication and time division multiple access (TDMA) methods. The whole inter-spacecraft communication chain is shown in Figure 32.
Figure 31: MASCOT communication subsystems overview (image credit: DLR, Telespazio VEGA)
• MASCOT Mobility Mechanism: It fulfils two functions, namely up-righting MASCOT into the correct attitude after landing, and providing a hopping capability to relocate MASCOT onto a different site. The motion is generated by driving an off-centered mass to provide the adequate momentum. The mass is driven by an electric motor, which is controlled by a dedicated, dual, cold-redundant electronic unit called MMC (Mobility Mechanism Controller). Each MMC branch is connected to the OBC via a UART RS422 line.
Figure 33: Sketch of separation & surface operation concept (image credit: DLR)
The OPS (Optical Position Sensors), designed and manufactured by Cosine Research B.V., Leiden, The Netherlands, is a small sensor (32:6 mm x 27 mm x 21:6 mm) with a mass of about 28 g. It consists of an infrared LED and an appropriate photodiode (Figure 34). The light omitted by the LED is reflected by any object in the FOV (Field of View) of the sensor. Nearer objects reflect more light and light from objects further away than 12 cm can not detected any more. This gives the opportunity to detect objects in proximity in a certain direction. In addition the LED is on/off modulated by the OBC and the signal from the photodiode is correlated with this modulation. With this method one can distinguish between reflected light from the LED and background illumination e.g. from sunlight. Figure 35 shows the output voltage of the OPS as function of the distance from the object.
Five of these OPS systems are used on MASCOT, mounted on 5 different sides of the cuboid. Only the instrument side is not equipped with such a sensor. This space is needed for the scientific instruments. As one can see in Figure 35, the output voltage decreases again for distances smaller than 10 mm. To prevent such small distances, the sensors are not mounted flat to the surface but shifted inside the body of MASCOT.
Figure 36: Photo of the Mascot EQM, mounted on the Hayabusa-2 Flight Model at JAXA (image credit: DLR, CNES)
MASCOT instruments: (MicrOmega, CAM, MAG, MARA)
The payload consists of a suite of instruments which fit into the payload compartment of the structure (Figure 37). Each instrument has its own electronic unit, which provides the signal conditioning to the instrument sensors/detectors and, where present, to the actuators, converts the acquired measurements into digital format and transmits them to the OBC via dedicated serial lines. The OBC is in charge of configuring the instruments, driving their acquisition sequence and performing data processing realizing instrument specific data acquisition modes. The instruments are not designed to be fault tolerant, except for the data lines, which are connected to both OBC main and redundant sections.
As an in situ Science Landing Package, MASCOT will augment the science capabilities of the Hayabusa-2 mission on three levels:
1) ‘Context Science' – By coordinated and complementary observations of the instruments on board MASCOT, on the main spacecraft and in combination with the lab analysis of the returned samples, MASCOT will ensure ground truth, down to the microscopic scale, for the acquired scientific results of the other mission elements. The goal is to obtain a cross-scale link with respect to surface and subsurface science , combining all three data sets to obtain a profound knowledge of the asteroid.
2) ‘Stand-alone Science' – MASCOT can perform unique investigations of major science importance including those that only a landed package can do, such as geophysics. Direct in situ measurements by the MASCOT instruments will also enable the analytical characterization of the elemental, isotopic and molecular (organic and mineral) composition of 1999JU3's surface and its near-surface material for samples in their natural state. Since 1999JU3 is a C-type asteroid, the astrobiological relevance of such measurements is significant. In addition, complementing visual documentation at the level of microscope scale would place the analytical results in context to the sample setting. MASCOT will investigate several locations on the asteroid to determine compositional and structural homogeneity and heterogeneity.
3) ‘Reconnaissance & Scouting' - By supporting Hayabusa-2's task of obtaining samples relevant to the top-level science objectives, MASCOT can serve as a ‘scouting vehicle' for assessing candidate sampling sites before the Hayabusa-2 samples are acquired; in other words, if enough lead time for MASCOT deployment is accounted for in mission operations, the final selection of the sampling sites can be guided by the results from the MASCOT instruments.
MicrOmega is a near infrared spectrometer and a hyperspectral infrared microscope for in situ mineralogical analyses of the ground, developed by IAS (Institut d'Astrophysique Spatiale, Orsay). The objective is to characterize the composition of surface samples at their grain scale. Supported by CNES, MicrOmega/MASCOT has been developed within the frame of the Pasteur payload of ExoMars' rover. It is also derived from a first MicrOmega model which was developed for Phobos Grunt. MicrOmega/MASCOT will enable the first in situ microscopic characterization of a C-type asteroid.
MicrOmega/MASCOT acquires monochromatic images of samples, 128 x 128 pixels of 20 µm2 each. The samples are illuminated by an AOTF-based dispersive system, onto a 2D HgCdTe array, cooled by a dedicated cryocooler. By scanning the illumination wavelength, over the spectral range 0.9 to 3.5 µm with steps of 20 cm-1, 3D (x,y,λ) image cubes are built, in about 10 mn.
The overall instrument, including its driving electronics, has a mass of 2 kg. The detector assembly, the illuminator and the analog electronic box of MicrOmega are shown in Figure 38.
The camera is provided by DLR (Figure 39) is mounted inside the lander slightly tilted, such that the center of its FOV (Field of View) is aimed at the surface at an angle of 22º with respect to the surface plane. This implies that both the surface close to the lander at a distance of 15 cm and the horizon are in the FOV. The optics are designed according to the Scheimpflug principle, which ensures that the entire scene is in focus.
This wide angle camera uses LED illumination for nighttime operation. It is equipped with a 1024 x 1024 pixel CMOS sensor sensitive in the 400-1000 nm wavelength range, and has a FOV of 60° x 60°. The instantaneous field of view is 2.1 mrad/pixel. The camera images will be high priority and involve five pictures per day and four per night. Images will also be made during descent and hopping.
Each photo has a volume of 14.7 Mbit. The instrument has heritage from ExoMars, Rosetta/Philae and the ISS.
MAG is provided by TU Braunschweig, Germany. As indicated by meteorite composition, the asteroid material is magnetic, but how is the magnetic field arranged? In order to answer the question about the magnetization state of asteroids, MASCOT MAG shall:
- Observe magnetic field profile during descent and hopping
- Determine global and local magnetization of the asteroid.
This will provide new data on asteroid formation and history, and provide deeper insight on early planetary formation stages. MAG (Figure 4) is a three-axial fluxgate vector compensated magnetometer based on a long heritage (Themis, Rosetta, VeX, BepiColombo).Its main performance parameters are :
- Dynamic range is ±50000 nT
- Sensor noise (@1Hz) 10 pT/√Hz
- Resolution 6 pT
- Sampling rate 10 Hz
Figure 40: Schematic view of MAG (image credit: TU Braunschweig)
MARA (MASCOT Radiometer):
MARA is a multispectral instrument, provided by DLR, to measure the radiative flux emitted from the asteroid's surface using thermopile sensors. Six individual filters will be employed to measure the flux in the wavelength bands between 5.5-7, 8-9.5, 9.5-11.5, 13.5-15.5, 8-14, and 5-100 µm. The primary scientific goal of the MARA instrument is the determination of the asteroid's thermal inertia, the secondary goal is the characterization of the surface mineralogy.
To determine the surface thermal inertia, MARA will measure the temperature of the asteroid's surface over the period of a full rotation using the long wavelength channel from 5 - 100 µm. In addition, the emissivity of the surface can be determined from the flux in the bandpass filters from 5.5-7, 8-9.5, 9.5-11.5, 13.5-15.5, and 8-14 µm. Thermal inertia can then be determined from an investigation of the surface radiative energy balance.
The mineralogy of the surface can be characterized from an investigation of the radiative flux in the same bandpass channels, as rock forming minerals like olivine and pyroxene have characteristic absorption features in the channels covered by MARA. In addition, the 8 to 14 µm filter is identical to the filter used by the thermal mapper on the Hayabusa-2 spacecraft, such that measurements by MARA can be directly compared to the results obtained from the spacecraft. In this way, MARA can provide ground truth at small scales, thus providing context for the spacecraft measurements.
Figure 41: Photo of the MARA sensor head flight model (image credit: DLR)
MARA is mounted in between the infrared spectrometer and the camera, such that the FOV of MARA and the camera overlap. MARA consists of the following functional elements:
- Sensor head
- MARA electronics
- In-flight calibration surface.
Overview of the MASCOT operational concept:
In the context of Hayabusa-2 mission, the following major flight operational phases are foreseen for MASCOT:
1) Cruise phase (~4 years): MASCOT is normally off controlled in temperature by Hayabusa-2. At agreed times, MASCOT is activated from the ground to perform checkouts and instrument calibrations; Special operations, like PRM activation, instrument commissioning, on-board software maintenance are also foreseen.
2) Asteroid approach and global mapping: This phase is to characterize the asteroid's surface and select the target landing site based on scientific relevance and lander's constraints; Operational preparation for landing will occur: Initial on-board parameter tuning, science command sequences upload, on-asteroid operations rehearsal.
3) Pre-separation: Final on-board parameter tuning, final science command sequences upload, battery depassivation, full MASCOT check-out, lander pre-heating.
4) SDL (Separation, Descent and Landing) and on-asteroid operations: During an Hayabusa-2 descent maneuver, MASCOT will be deployed towards the asteroid's surface. Primary batteries will supply the power for the mission. Due to the short lifetime (10-16 hrs depending on the mission profile) all operations for science data collection, processing, downlink to Hayabusa-2, uprighting and hopping maneuvers will be fully autonomous, driven by the MAM (Mission Autonomy Manager), an OBSW package. This phase will last up to EOL (End-Of-Life), corresponding to the discharge of the primary batteries.
Autonomous Operations during SDL and On-Asteroid Phase:
After the global mapping and pre-separation phase operations, MASCOT will be released towards the asteroid's surface: Autonomous operations driven by the MAM will start at this point, after the separation detection. Manual commanding from the ground is still possible in parallel to MAM, but its use is only in case of unforeseen contingencies, not handled on-board. MAM commanding can be paused if necessary to avoid interference in the commanding.
MASCOT deployment will occur either during a dedicated Hayabusa-2 descent or during one of the Hayabusa-2 sampling touchdown rehearsals. This maneuver foresees Hayabusa-2 to descent to a separation altitude less than 100 m at which point MASCOT will get ejected via a spring mechanism. As MASCOT has no propulsion system, its descent to the surface of the asteroid will be completely passive, under the effects of the weak asteroid gravity field, followed by a bouncing period, before it comes to rest in
During descent, MASCOT will measure the asteroid magnetic field by the MAG instrument. The MAM will also schedule several image takings of the asteroid by the CAM instrument (a specific CAM image processing algorithm is implemented in the OBSW to discard dark - night sky - or saturated - sun facing - pictures).
During descent, MASCOT will measure the asteroid magnetic field by the MAG instrument. The MAM will also schedule several image takings of the asteroid by the CAM instrument (a specific CAM image processing algorithm is implemented in the OBSW to discard dark - night sky - or saturated - sun facing - pictures).
Once at rest on the surface, the MAM will start an upright maneuver, in case MASCOT is not laying with the -Z side (bottom) in contact with the surface: This is necessary for CAM, MARA and MicrOmega instrument science operations to characterize the asteroid soil. The upright maneuver may take several attempts and a specific FDIR algorithm has been devised to maximize the chance of success, to cope against adverse surface morphology and possible on-board failures.
Landing is currently foreseen at around the asteroid midday, so the MAM will schedule a day and a night science measurement cycle on the first landing location. Afterwards, the mobility mechanism will be activated by the MAM to send MASCOT in an uncontrolled hop across the asteroid's surface at a varying distance of up to a few 10 m. Further scientific activities will take place, and then, depending on the power and data download status, a second hop will be considered.
After MASCOT's release, Hayabusa-2 will return to its home position (HP), hovering 20 km above the asteroid's surface (Figure 42). During the ascent phase, pictures will be taken from Hayabusa-2 of the MASCOT landing site, using the asteroid camera.
The selected landing site will have to ensure the following characteristics:
- daylight duration between 50% and 70% of the asteroid's rotation period, due to thermal and scientific reasons (one asteroid day is estimated to be ~7.6 hours)
- the duration per asteroid rotation period of visibility of the landing site from Home Position must be over 40% (TM/TC link constraint)
- the velocity at touchdown must be smaller than half the escape velocity
- average surface temperature between -50ºC and +25ºC (under re-assessment) to grant acceptable landing conditions.
SRC (Sample Return Capsule)
The SRC is designed to deliver the asteroid samples to Earth. A series of major TCM (Trajectory Correction Maneuvers) are to be carried out starting a month before the reentry (R) day. The spacecraft is to be inserted into the reentry orbit targeting the landing center at the final TCM on about R-3 day. A schematic view of the reentry sequence is shown in Figure 43. - The SRC is separated from the mothership about 8 hours before the reentry and the spacecraft will enter the escape trajectory after orbital maneuvering by the RCS while the SRC enters Earth's atmosphere at a velocity of about 12 km/s. After passing through the severe aerodynamic environment, SRC deploys the parachute at the altitude of about 5 km and it descends slowly with beacon signal emitted for localization until landing on the ground. 47)
Figure 44 shows the aerodynamic heating environment of the Hayabusa-2 SRC when compared with that of Hayabusa-1 as a function of the flight path angle and the entry velocity. The main parameters dominating the reentry flight environment are the flight path angle and the entry velocity while the effect of the entry position (latitude/longitude) and the flight azimuth direction is secondary, so that we can neglect them at the present discussion. A maximum heat flux of 13 MW/m2 is predicted on the stagnation heat transfer rate with total heat input of 270 MJ/m2.
Design and verification tests for the Hayabusa-2 SRC:
There are strong demands of short-term development and cost reduction as well as improvement of the reliability for the Hayabusa-2 project. Therefore, the basic design concept for the Hayabusa-2 SRC is not to change the design from the former SRC except for the following items:
1) During the period of over 10 years since the development of the Hayabusa-1 SRC, some components have gone out of production. Since the many electronics parts, including FPGAs, used as the core processors of the SRC instrument module are discontinued, the onboard electronics of the instrument module need to be redesigned.
2) The SRC design needs to be improved based upon the experience of Hayabusa-1 operation. Examples of this case include employing a new HK telemetry for monitoring of the on-board clock accuracy and a backup timer for generating the parachute deployment trigger for the case of an acceleration sensor malfunction.
3) The items judged by their design shall be changed to be compatible with the change of the design condition such as environmental requirements and the design standard of Hayabusa-2 system. Examples of this case include a change of the criterion on the tensional force of the Marman Clamp band of the SRC separation mechanism.
4) The items judged by their design need to be improved to ensure higher reliability based on the result of the design review including the evaluation of the reliability analysis such as FMEA, FTA and SPFA. Examples of this case include a change of an ignition circuit design, strengthening of the puller used for the SRC separation mechanism, and an adoption of the additional O-ring to prevent the inflow of high-temperature gas during reentry operation.
Overview of the SRC Subsystem:
Hayabusa-2 SRC is designed with the following functions in its compact body:
• Sampler Interface: SRC has an interface with the sampler-container system for collecting and storing asteroid sample. The sampler-container is fixed to the inside of the SRC at the time of the launch. The sampler-catcher is held in the sampler-mechanism. After Hayabusa-2 approaches the target asteroid and collects a "sample", the sampler-catcher is conveyed into the sampler container and is fixed to it with the latch mechanism. The SRC conducts the Earth reentry in the configuration of Figure 45.
• Mothership Interface / SRC Separation Mechanism: In addition to its mechanical I/F, the SRC has electric and thermal I/F with the mother ship, such as HK telemetry, command, heater control, and external power supply from the mothership. The SRC separation mechanism holds the SRC on the mothership during the orbital flight. The mechanism deploys the Merman Clamp band and ejects the SRC from the mothership with proper velocity and spin rate given by the helical coil spring.
• Heatshield and Thermal Protection Subsystem: In order to protect the internal instruments from the aerodynamic heating environment during reentry operation, the SRC is wrapped with an ablator made of carbon phenolic resin. The ablator consists of four parts; the fore body ablator, the aft ablator, the support ablator and the sampler ablator. With the aim of avoiding soakback from the ablator, the instrument box in the SRC throws the fore body ablator and the aft ablator away after passing through the excessively severe aerodynamic environment (Figure 46).
• Descending Subsystem: After the separation of the ablators, the instrument box connected to a parachute starts slow descending. The descending system consists of the cross-type parachute, the parachute deployment mechanism and the ignition circuit (IG-BOX). The instrument box is also equipped with the parachute anchor release mechanism in order to avoid breakage due to dragging by strong ground-winds after landing.
• Beacon Transmission: The beacon antenna is pulled out and hanged from the bottom of the instrument module (I/M) by a small mass at the time of the parachute deployment. And simultaneously, I/M starts transmitting the beacon signal for the localization from the ground stations. The beacon is transmitted even after the instrument box lands on the ground according to the prescribed sequence.
• Reentry Flight Measurement: Hayabusa-2 has a new capability to measure acceleration, attitude motion (angular rate) and the temperature of each part inside the SRC during reentry operation. The measurement data is stored to a memory and is retrieved after recovery. This function is implemented in a new instrument: REMM (Reentry Flight Measurement Module).
REMM is an instrument which was newly developed to add the function of flight environment measurement to SRC. REMM measures the heat shield temperatures, the acceleration and the attitude motion of the SRC during the reentry and records the measurement data to a non-volatile memory. Since there is no structural design change for Hayabusa-2 SRC, only very limited space and mass can be allocated for the REMM. A COTS based design was adopted in design of REMM to meet the requirement under this very strict constraints.
Table 6: Specification of REMM
Figure 47: REMM functional block diagram (image credit: JAXA)
Figure 48: REMM FM (Flight Model) and its configuration in SRC (image credit: JAXA)
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: "Observation of the Earth and Its Environment: Survey of Missions and Sensors" (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates (email@example.com).