Hayabusa2 Information
Fact Sheet
Edition time: Immediately after
arrival at the asteroid
Ver. 2. 3
2018. 07. 05
Hayabusa2 Project TeamTable of Contents
1. tOverview
2. tThe spacecraft
3. tHistory of the mission
4. tOrbit
5. tNear-asteroid operations
6. tOperations
7. tTarget body
8. tScience
9. tInternational cooperation

3
14
54
77
90
97
104
122
130

&

1. Overview'Overview of Hayabusa2
ObjectiveWe will explore and sample the C-type asteroid Ryugu, which is a more primitive
type than the S-type asteroid Itokawa that Hayabusa explored, and elucidate
interactions between minerals, water, and organic matter in the primitive solar
system. By doing so, we will learn about the origin and evolution of Earth, theoceans, and life, and maintain and develop the technologies for deep-space return
exploration (as demonstrated with Hayabusa), a field in which Japan leads theworld.

Expected results and effects
By exploring a C-type asteroid, which is rich in water and organic materials, wewill clarify interactions between the building blocks of Earth and the evolution of
its oceans and life, thereby developing solar system science. Japan will further its worldwide lead in this field by taking on the new challenge ofobtaining samples from a crater produced by an impacting device.
We will establish stable technologies for return exploration of solar-system bodies.

Features:World’s first sample return mission to a C-type asteroid.
World’s first attempt at a rendezvous with an asteroid and performance ofobservation before and after projectile impact from an impactor.
Comparison with results from Hayabusa will allow deeper understanding of the
distribution, origins, and evolution of materials in the solar system.

International positioningªJapan is a leader in the field of primitive body exploration, and visiting a type-C
asteroid marks a new accomplishment.
This mission builds on the originality and successes of the Hayabusa mission. In
addition to developing planetary science and solar system exploration
technologies in Japan, this mission develops new frontiers in exploration ofprimitive heavenly bodies.
NASA too is conducting an asteroid sample return mission, OSIRIS-REx (launch:
2016; asteroid arrival: 2018; Earth return: 2023). We will exchange samples andotherwise promote scientific exchange, and expect further scientific findings
through comparison and investigation of the results from both missions.

(Illustration: Akihiro Ikeshita)
Hayabusa 2 primary specifications
Mass
Approx. 609 kg
Launch3 Dec 2014
MissionAsteroid return
Arrival27 June 2018
Earth return2020
Stay at asteroid
Approx. 18 months
Target body
Near-Earth asteroid Ryugu
Primary instruments
Sampling mechanism, re-entry capsule, optical cameras, laser range-finder,
scientific observation equipment (near-infrared, thermal infrared), impactor,
miniature rovers.

(Mission significance
1. Scientific significance“Where did we come from?”— Origins and evolution of the solar system and thebuilding blocks of life
The materials that formed the Earth, its oceans, and life were present in the primordial cloud from which oursolar system formed. In the early solar system, these materials were in contact and able to chemically interact
within the same parent objects. These interactions are retained even today in primitive bodies (C-type
asteroids), so returning samples from these bodies for analysis will elucidate the origins and evolution of thesolar system and the building blocks of life.

2. Technical significance“World-leading technology”— Continuance and development of Japan’s unique deepspace exploration technologies
As the world-first asteroid sample return mission, the Hayabusa mission incorporated a variety of newtechnologies. Continuing that experience, we will establish technologies that allow more reliable deep space
exploration. Taking on these new technical challenges will create new opportunities for the future.

3. Exploration significancett“Exploring new frontiers”—Effects including scientific innovation, contributions toindustry and society, improved international presence, youth development
By entering these unexplored fields, we will create new scientific technologies, contribute to industry, andfurthermore contribute to society by providing knowledge related to the issue of Earth-threatening asteroids,
space resource utilization, and targets for manned exploration.

)Mission objectives
 Scientific objective 1: Solving mysteries related to the processes of material

evolution in the solar system
We will investigate a type-C asteroid from a materials science perspective. In
particular, we will elucidate interactions between minerals, water, and organic
materials.

 Scientific objective 2: Solving mysteries related to the evolutionary process ofplanets
We will examine the formation process of asteroids through direct study of theintegration of materials into asteroids, their internal structure, and subsurface
material.

 Engineering objective 1: Establishment of technologies for deep space sample

return exploration
We will bring these technologies to maturity by improving their robustness, stability,
and operability.

 Engineering objective 2: Demonstration of space impactor technology
We will demonstrate collision of an impactor on a celestial body.

*Mission flow
Launch3 Dec 2014

Arrival at asteroid
27 Jun 2018t▲Earth swing-by
3 Dec 2015

Examine the asteroid by remote sensing
observations. Next, release a small lander androver and also obtain samples from the surface.

Earth return
Late 2020

Depart asteroid
Nov–Dec 2019

Create artificial
craterSample analysis(Illustrations: Akihiro Ikeshita)

After confirming safety, touchdown
within the crater and obtain
subsurface samples

Releaseimpactor

Use an impactor to create an
artificial crater on theasteroid’s surface+Planned operationsV Depart Earth
V IES trial
V Begin IES-poweredflightV Earth swing-by
V Subsequent long-term
IES operation

V Interim operations

V Impactor operations(crater creation)
V Debris/ejecta avoidance
operations(Illustrations: Akihiro Ikeshita)

V Touchdown in artificial
craterV Asteroid rendezvous by
optical navigation

V Maintain position
V Global mapping of
the asteroid by
V Landing practice & implementation proximity
observations
V Lander/rover separation
V Touchdown and sampling

V Depart asteroid

V Earth re-entry,Overall schedule
2015
3

12

Event

Initialoperation.

2016
10 12ESA (MLG/WLH）
test operations(21‒22 May)Mar

Operations for travel to asteroid

Earth swing-by(3 Dec)Launch(3 Dec)Ion engine operations

2018

4

SwingbyEDVEGA

2017

Jun

May

Nov

12
Earth return

Asteroid proximity operations

Depart Ryugu(Nov‒Dec)

Re-e12ntry

Capsule re-entry(late 2020)

Interim period
Optical navigation(Solar conjunction)
May Jul Nov Dec

May
Mar

2020

6 7
Approach
Ryugu arrival(27 Jun)Southern hemisphere station
operations (CAN/MLG)
Oct

2019

Apr

Jan

Jun

TBD

TBD TBD

TBD-Preliminary asteroid approach schedule
Year

Month/Day

Item

Status2018

10 JanPhase 3 ion engine operations begin

Complete

3 Jun

Ion engine operation ends

Complete

3 JunStart of asteroid approach (dist. 3,100 km)

Complete

27 JunArrive at asteroid (alt. 20 km)

Complete

Late Jul

Medium altitude observation #1 (alt. 5 km）

Est.

Aug

Gravity measurement descent (alt. 1 km）

Est.

Late Aug

Decision of landing sites

Est.

Sep‒Oct

Touchdown operation slot #1

Est.

Sep‒Oct

Rover descent operation slot #1

Est.

Nov‒Dec

Interim operations (communication unavailable）

Est.

Jan

Medium altitude observation #2 (alt. 5 km）

Est.

Feb

Touchdown operation slot #2

Est.

Mar‒AprCrater creation operations

Est.

Apr‒May

Touchdown operation slot #3

Est.

Jul

Rover descent operation slot #2

Est.

Aug‒Nov

Stay in asteroid vicinity

Est.

Nov‒Dec

Depart asteroid

Est.

2019

Note that the above schedule is subject to change according to various factors after arrival at Ryugu;
all dates are tentative, other than those marked “complete.”%$Initial version

Mission patch
From preparation through launch, the JAXA SPace
Exploration Center (JSPEC, the program group for lunar andplanetary exploration) primarily led the Hayabusa2 project.
The patch shows touchdown on the target asteroid 1999 JU3.
The path past the Earth, Moon, and Mars indicates our intent
to advance technologies and science for future solar system
exploration. The shape of the red background indicates thetwo high-gain antennas that are characteristic of Hayabusa2.

Version from Earth swing-by to asteroid arrival
With the 3 Dec 2015 Earth swing-by and start of theapproach toward Ryugu (the asteroid was named in Sept.
2015), we changed the background colors to blue, indicating
departure from the vicinity of Earth and plunging deeper into
space. This works well with imagery from the Japanese folk
tale from which Ryugu gets its name: Urashima Taro rode ona sea turtle’s back to Ryugu Palace, deep beneath theocean.

(© JAXA) 	

%%

Mission patch
Version from asteroid arrival
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(© JAXA) 	

%&

Asteroid exploration at JAXA/ISAS
Hayabusa2

Hayabusa
2003–20102014

Subsequent missions

2011–2020

We are considering
exploration of the
Jovian Trojans bysolar sail

To more primitive bodies

S-typeC-typeTo more advanced technologies

D-typeTo farther distances

Asteroid belt(© JAXA) 	

Note: Ryugu is a C-type asteroid, exceptional in that it exists
near the orbits of Earth and Mars.

13

History of asteroid exploration planning
Concept fromaround1985

Significant
problemoccurs

Year

Today (Jun 2018)

00 01 02 03 04 05 06 07 08 09 10 11 12 13 14 15 16 17 18 19

MUSES-CDevelopment
since 1996

▲LaunchHayabusa2 launcht▲Project

Hayabusa2 investigation

Around2000

PostMUSES-C

Sample analysis

Hayabusa operation

PostHayabusa

Hayabusa Mk2New mission

Marco Polo▲Apr 2004ª Launch of small body
exploration working group

Hayabusa 2 reviews, etc.
▲

2006ªStart
investigation of
Hayabusa2 mission▲Jun 2007ªLaunch of
Hayabusa2 prep. teamStarted as a “revenge” mission for
Hayabusa, initially as a “copy”
spacecraft; target launch of 2010

▲May 2011ª Launch of
Hayabusa2 project team2006.10–11 ª Mission definition review (MDR), system
requirements review (SRR), system definition review (SDR)→
2010 launch2009.6-7ª ΔMDR (mission scope expansion)→ 2014 launch2009.12ª SRR
2010.8ª Space Development Committee pre-evaluation #1
2011.3ª SDR
2011.4–9ª Preliminary design review (PDR)
2011.6–2012.1ª Space Development Committee pre-evaluation #2
2012.3tª Critical design review (CDR)

%(

2. The spacecraft%)Primary spacecraft components
Deployable camera(DCAM3)

X-band highgain antenna

X-band low-gain
antennaX-band mediumgain antenna

Solar array paddle

Ka-band highgain antennaOptical navigation Laser altimeter
cameraLIDAR
ONC-T

Near-infrared
spectrometer
NIRS3

Thermal infrared
cameraTIR

Scientific observation equipment
Star trackers

Ion engineNear-infrared spectrometer(NIRS3)

RCS thrusters
Optical navigation
cameraONC-T, ONC-W1

Re-entry capsuleSampler horn

Laser altimeter(LIDAR)

Optical navigation
camera ONC-W2

Small lander & rover
MASCOT

Minerva ©

MASCOT lander
Thermal infrared camera(TIR)

II-1A

II-1B

II-2

Created by DLR andCNES
II-2: By Tohoku Univ. & the Minerva-II
II-1: By the JAXA Minerva-II teamConsortium

(© JAXA) 

Minerva-II rover

Small carry-on impactor (SCI)
Target markers × 5Sizeª 1 ×1. 6×1. 25 m (main body)t
ttttttSolar paddle deployed width 6 m
Mass ª 609 kg (incl. fuel)

%*Device names (1/2¨Deployable	
  Camera

X-­‐band	
  	
  Low-­‐gain	  Antenna	  X-­‐band	
  High-­‐Gain	
  Antenna	  X-­‐band	
  Medium-­‐gain	  Antenna	  

Solar	
  Array	  Paddle

Ka-­‐band	
  High-­‐gain	  Antenna

Star	
  trackers
Near-­‐infrared	
  Spectrometer	
  (NIRS3)

+Z

Re-­‐entry	
  CapsuleSampler	
  Horn(© JAXA) 	

Laser	
  Al�meter(LIDER)

Op�cal	  Naviga�on	
  Camera-­‐	
  Wide	  (ONC-­‐W2)

-­‐Y

-­‐X

17

Device names (2/2¨Ion	
  Engine

RCS	
  Thrusters	  (12)
Op�cal	  Naviga�on	
  	
  Camera–Telescopic	
  &	  Wide(ONC-­‐T,	
  ONC-­‐W1)

MASCOT	
  (Small	  Lander	  by	
  DLR/CNES)
	
  MINERVA-­‐II	
  (Small	
  Rovers)

+Z

Thermal	  Infrared	
  Imager	
  (TIR)
Small	  Carry-­‐on	  Impactor	
  (SCI)

+Y
+X

Target	  Markers	
  (5)

(© JAXA) 	

%,Spacecraft dimensions
188

136
160

125
100
600

423

40
100

Numbers show approximate
length in cm(© JAXA) 	

%-Primary engagement testing(© JAXA) 	

26 Dec 2012: JAXA Sagamihara Campus&$Completion of primary engagement testingJun 2013ª JAXA Sagamihara Campus
(© JAXA) 	

&%Flight model(© JAXA) 	

31 Aug 2014: JAXA Sagamihara Campus&&Comparison of Hayabusa and Hayabusa2 (1/3)
Hayabusa

Hayabusa2 Dimensionsª Approx. 1 × 1.6 × 1. 25µ(main body)
Mass: 600 kg (with fuel)

Dimensionsª Approx. 1 × 1.6 × 1. 1 µ(main body)
Mass: 510 kg (with fuel)

X-band high-gain antenna	

XTop view
X-band high-gain antenna	

①

w

Solar array paddle

Ka-band high-gain antenna	

Ka

w

③	


Small lander

④	

Improvements over Hayabusa

②	xzy
Ion engine(© JAXA) 	

`Communication system: A new Ka-band communication system was added for high-speed communication. The high-gain antenna wasmade into a planar antenna.
aIon engine: Improved durability, stronger propulsionb Mobile Asteroid Surface Scout (MASCOT) small landerª Developed in Germany and France for landing and data acquisition on an
asteroid surface.
cAttitude control device (reaction wheel): Two of the three installed on Hayabusa malfunctioned, so four are mounted on Hayabusa2 andfurther measures have been taken to avoid problems.

&'Comparison of Hayabusa and Hayabusa2 (2/3)
Hayabusa

Hayabusa2 

Bottom view

⑦	

Near-infrared
spectrometer

Re-entry capsule⑥	

Small rover
Thermal infrared
camera⑧	

 Target marker
Impactor

Laser altimeter
altimeter
Laser
Sampling device
Optical navigation camera(© JAXA) 	

⑤	

Improvements over Hayabusa
dImpactorª A new device for creating an artificial crater on the asteroid surface, then collecting subsurface materials.
eRCS thrustersª Improved propellant plumbing as a countermeasure against the malfunctions on Hayabusa and Akatsuki.
fMission equipment: Newly developed and improved equipment for exploration of a type-C asteroid.
g Target markers: Increased from three on Hayabusa to five on Hayabusa2 to realize a pinpoint landing.

&(Comparison of Hayabusa and Hayabusa2 (3/3)

Hayabusa

Hayabusa	
  2	
  

Main	
  body	  dimensions	
  

1	
  	
  ×	
  1.6	  	
  ×	
  1.1	
  µ

1	
  	
  ×	
  1.6	  ×	
  1.25	
  µMass	
  (with	  fuel)

510	
  ´³

609	
  ´³

Launch	  year	  and	  rocket9	
  May	
  2003,	  M-­‐V-­‐5	
  rocket3	
  Dec	  2014,	  H-­‐IIA	
  rocket	
  ﬂight	
  26

Communica�ons	
  frequencies

X-­‐band	
  (7–8	
  GHz)

X-­‐band	
  (7–8	
  GHz)uKa-­‐band	
  (32	  GHz)

Near-­‐infrared	
  spectrometer,	
  ﬂuorescent	  X-­‐ray	  spectrometer,	
  mul�band	
  Mission	
  equipment
spectroscopic	
  camera,	
  laser	  al�meter,	  MINERVA,	
  samplerNear-­‐infrared	
  spectrometer,	
  thermal	
  infrared	
  camera,	  op�cal	
  naviga�on	
  camera,	
  laser	  al�meter,	
  MINERVA-­‐II,	
  MASCOT,	
  impactor,	  separa�on	
  camera,	
  sampling	
  device

Explora�on	
  period Approx.	
  3	
  months

Approx.	
  18	
  months	
  (planned)

Samples2	
  (surface	
  only)

3	
  (surface	
  and	  a�empted	
  subsurface)

Earth	
  return

13	
  Jun	  2010Late	
  2020	
  (planned¨&)List of mission equipment
DeviceRoleOptical Navigation Camera(ONC)

Telescopic and wide-angle cameras centered on visible wavelengths, withrespective viewing angles of 6 and 60 deg. These are used for scientificobservations and navigation.

Near-infrared spectrometer(NIRS3)

Performs spectroscopic observations of near-infrared rays including the 3micron band. The viewing angle is about 0.1 deg.

Thermal infrared spectrometer(TIR)

Images the asteroid at mid-infrared ranges including the 10-micron band.
Viewing angle is a little over 10 deg. 

Laser altimeter(LIDAR)

Measures the distance between the probe and the asteroid surface. Also
acquires scientific data such as asteroid topography, gravity, and albedo.
Measurement ranges are 30 m–25 km.Sampling device(SMP)

Acquires samples from the asteroid surface.
Slight improvements over the Hayabusa sampling device.

Impactor(SCI)

Accelerates a 2-kg copper mass to 2 km/s to collide with the asteroid surface,forming an artificial crater.

Deployable camera (DCAM)

Separates from the spacecraft to image the impactor operation.

Small rovers(MINERVA-II-1 (A, B), 2)

Descends to the asteroid surface for investigations. Three rovers similar to
MINERVA mounted on Hayabusa. 

Small lander(MASCOT)

Descends to the asteroid surface to acquire data through four observation
devices. Created by DLR (Germany) and CNES (France).
Observation devices: MicrOmega, MAG, CAM, MARA&*

Remote sensing equipment
Optical Navigation Camera (ONC¨

Thermal Infrared Camera (TIR¨

ONC-T (telescopic)ttONC-W1,W2 (wide-angle)
Imaging for scientific observation and navigation

Near-infrared Spectrometer (NIRS3¨

Infrared spectra including the 3-µm band: investigates
mineral distributions on the asteroid surface(© JAXA) 	

8–12 µm imaging: Measures asteroid surface temperature

Laser Altimeter (LIDAR¨

Measures distance between the asteroid and the spacecraft
in a range of 30 m–25 km&+Optical navigation camera (ONC¨
ONCªOptical Navigation Camera
Objective: Images fixed stars and the target
asteroid for spacecraft guidance and scientificmeasurements
Scientific measurementsª
V Form and motion of the asteroid:
Diameter, volume, direction of inertial principal axis,
nutationV

Global observations of surface topography
Craters, structural topography, rubble, regolith
distribution

V Global observations of spectroscopic properties of
surface materials
Hydrous mineral distribution, distribution of organic
matter, degree of space weathering
V

High-resolution imaging near the sampling point
Size, form, degree of bonding, and heterogeneity ofsurface particles; observation of sampler projectiles
and surface markings

V Elucidation of features of target
asteroid
V Distribution of hydrous minerals andorganic matter, space weathering,boulders
V Sampling site selection
V Basic information on where to collectasteroid samples
V Ascertaining sample state
V High-resolution imaging of samplingsites

(© JAXA) 

ONC-T
Detector

W1
FoV
T
FoV

ONC-W1

ONC-W2

2D Si-CCD (1024 × 1024 px)	Viewing
direction

Downward (telescopic)

Viewingangle

6.35° × 6.35°�

65.24° × 65.24°�

100 m–∞

1 m–∞

Focal length

Downward (wideangle)

Sideward (wideangle)

Spatialresolution

1 m/px @ 10-km alt.
1 cm/px @100-m alt.

10 m/px @10-km alt.
1 mm/px @1-m alt.

Observation
wavelength

390, 480, 550, 700, 860,
950, 589.5 nm, and wide �

485–655 nm&,Laser altimeter (LIDAR¨
LIDAR: Light Detection And RangingPulse-type laser altimeter A pulse YAG laser with a 1.064-µm
wavelength is emitted toward the target
object, and the altitude is measured bymeasuring the return time of the laser
beam.
 The LIDAR aboard Hayabusa 2 could
perform measurements from 30 m–25 km.
 LIDAR is a navigation sensor used forapproach and landing at a target, and ascientific observation device used tomeasure shape, gravity, and surface
characteristics, and for dust observations.
 It also has a transponder function that canperform space laser ranging (SLR)
experiments with ground LIDAR stations.

(© JAXA) 

Short-range telescopeBeam expander for illumination

Long-range Cassegrain telescope
Laser altimeter engineering model
Scientific objectives
V Terrain and gravity field observations of thetarget asteroid
V Observations of albedo distribution at various
surface points
V Observations of dust floating around the asteroid

V Asteroid form, mass, porosity, and deviation
V Asteroid surface roughness
V Dust floating phenomena&-Near-infrared spectrometer (NIRS3)
NIRS3: Near-infrared Spectrometer(‘3’ from 3 µm)Infrared absorption of hydroxyl groups and watermolecules is observed in 3-µm band reflection
spectra in the near-infrared region. NIRS3
investigates distributions of hydrous minerals on theasteroid surface by measuring reflection spectra in
the 3-µm band.
OH H2O

Observation wavelength range: 1.8–3.2 µm
Wavelength resolution: 20 nm
Full field of view: 0.1 deg
Spatial resolution: 35 m (20-km alt.)
ttttttt
2 m (1-km alt.¨V Detector temperature: –85 to –70 °C
V S/N ratio: 50+ (wavelength 2.6 µm)

Orgueil (CI)

Absorption form anddepth vary dependingon the type andamount of hydrous
mineral

1.5
Scaled reflectanceReflectance
Relative

V
V
V
V

Murchison (CM)1.0

Renazzo (CR)Near-infrared reflection
spectra of carbonaceouschondrites

0.5

Allende (CV)0.0(© JAXA) 	

2.0

2.5
Wavelength (µm)

3.0Wavelength (µm)

'$

Thermal infrared camera (TIR¨
TIR=Thermal Infrared Imager
The surface temperature of the asteroid changes
over the day, rising in sunlight and decreasing atnight.
Diurnal change in surface temperature is large infine soils like sand and highly porous rock, andsmall in dense rock.
We will examine the physical state of the asteroid’s
surface by 2D imaging (thermography) of thermalradiation from the asteroid.

Large

400	
  
350	
  

Temp (K)
Temperature	
  [K]: 2D uncooled bolometer: 8–12 µm: –40 to 150 °C: 0.3 °C
: 328 × 248 (effective): 16°×12°: 20 m (20-km alt.¨
5 cm (50-m alt.¨

Stone

GravelSand Detector
 Observation wavelength Observed temperatures
 Relative accuracy Dimensions
 Viewing angle
 Resolution
tttttttt

Small

Temperature change

Night

Night

Day

I=10

300	
  

I=50
I=100

250	
  

I=200

200	
  

I=500

150	
  

(© JAXA) 	

100	
  
-180

I=1000

I=2000
-135

-90

-45

0

45

Longitude	
  [deg]tLongitude

90

135

180

'%

Sampler (SMP¨
Sampler horn

V Device for acquiring samples from the asteroid surface
V The basic design is the same as that for Hayabusa. As soon as
the tip of the cylindrical horn touches the asteroid surface, asmall projectile is shot from within the horn and rising surface
ejecta are caught in a catcher in the upper part of the horn.
V Sealing performance is improved in Hayabusa2; a newly
developed metal seal system ensures that volatile gases canbe brought back securely. Noble gases can also be collected.
V The sample catcher has been improved over that onboard
Hayabusa, and now contains three chambers instead of two.
V As a further improvement in Hayabusa2, there are small folded
parts on the tip of the horn, as shown in the figure. Grains of 1–
5 mm are caught in these folds, and the catcher is designed sothat samples continue rising when the spacecraft suddenly
ceases its ascent, thereby entering the catcher. This provides a
backup for sampling by projectile.

(© JAXA) 

Grains loaded
here

'&

Impactor (SCI)
SCIªSmall Carry-on Impactor■ Objectiveªl We will investigate the internal structure of the asteroid through surface changes before and after
projectile impact. We will also conduct remote observations of the exposed subsurface material toinvestigate physical properties of the surface.
l We also perform sampling from craters formed by projectiles, collect “fresh” substances from beneaththe surface, and investigate differences from surface materials.
l We will perform “space impact experiments” on actual asteroids to obtain data necessary for celestial
collision science.
■ Crater creation: Impact with a high-speed projectilel Can be performed with a mounted small, lightweight device.
l Results in less soil contamination than methods using explosives to expose asteroid surface
materials.
l Impacting projectiles are made from pure copper so that they can easily be distinguished from!
substances present in asteroids.
■SCI technologyl Application of technologies for molding explosive
chargesl Accelerates a 2 kg copper liner to approximately 2
km/s within approximately 1 ms(© JAXA) 	

!
!
!



Copper plate (liner) deforms during flight''Impactor: Structure

Direction of liner ejection	

External appearance

Cross-section

Copperplate

Explosive part

Accelerates the metal liner part (explosives attached to metal casing)u

Form: Cylindrical (diameter 265 mm)uLiner (becomes projectile): Pure copperu

Explosive: HMX-type PBX (plastic bonded explosive)u

Mass: Approx. 9.5 kg (explosive: 4.7kg, liner: 2.5 kg)uLiner thickness: Approx. 5 mm

Prototype(© JAXA) 	

'(

Impactor testing
Testsª We performed actual-use tests of the impactor to obtain technical data related to the projectile’s
speed, form, and attitude (17–27 Oct 2011).
Resultsª We obtained data from half- and full-scale models, which confirmed that the projectile forms a bell
shape in the explosive propulsion, rather than disintegrating.

Testing scene: Testing with a full-scale model.` Testing scene (moment of detonation)
The ignition point is surrounded by a 3-m concrete wall (right), withthe projectile fired to the lower left.

a Projectile form
The projectile travelling at approximately 2
km/sec. The outer diameter isapproximately 135 mm, and the mass isapproximately 2 kg. It takes a bell-like
form.

(© JAXA) 	

b Pierced targets
Targets pierced by the projectile. The interim targets trace the path
the projectile took to the final 4 x 4 m target, 100 m from theimpactor. c Moment of impact
The moment of impact into a dirttarget approximately 100 m fromthe firing position. (This is therear view of the target in image
3).

35

Deployable camera (DCAM3¨
DCAM3 = Deployable Camera 3

Successor to DCAM 1 and 2, mounted on the solar sail IKAROS.
This is a deployable camera for imaging the Small Carry-on
Impactor (SCI) and the asteroid as the projectile from the SCI hits
the surface, during which time the spacecraft will be sheltering ina safe zone. Imaging data are wirelessly transmitted to thespacecraft in real time.
n Engineering objectiveª Confirmation of impactor operations
V The spacecraft will be sheltering in a safe zone before SCI operation,and therefore has no way of confirming successful operations. SCI
operations are thus confirmed by releasing a deployable camera before
the spacecraft takes shelter and wirelessly transmitting acquired image
data.

DCAM3-Alens

DigitalantennaDCAM3-Dlens

Analog antenna

⌀ 78 mm × L 78 mm

SCI

FoV (74°¨–250 m

DCAM3

–1000 mImpact point

1999JU3(Diameter:
approx. 1 km )

n Scientific objective: On-site impact observation
V Continuous imaging of ejecta discharge will clarify relations between
asteroid surface conditions and ejecta emission phenomena.
Spacecraft takes
V We aim to identify the ignition point and the impact point of the impactor.
shelterV The produced ejecta will clarify crater formation processes on the
V While taking shelter, the spacecraft deploys the camera
asteroid.
at a position allowing views of the impact point from theside.
V The camera is separated so that its optical axis faces
the asteroid, and its mechanism separates while rotating
about the optical axis to stabilize its attitude.

(© JAXA) 	

'*Deployable camera (DCAM3¨n Overview of specifications and operationalplans
V Excluding lenses and the antenna, theseparable camera is a ⌀ 78 mm × H 81 mm
cylinder.
V Two cameras are mounted: an analog
camera that has low resolution but is capableof sending images in real time, and a digital
camera for digitally transmitting highresolution images.
V The image transmitter and the transmission
antenna are equipped with both analog anddigital systems.
V Batteries have relatively large capacities,
allowing for imaging and wireless datatransmission of up to 3 h (depending on
conditions).
V Image transmission is possible from up to 10
km from the spacecraft.

(© JAXA) 	

Analog camera
optics system

Separable camera unit

Analog imaging
transmission antenna

Digital camera
optics system

Separation
mechanism

Digital imaging transmission antenna'+MINERVA-II
V Three robots will travel across and explore theasteroid surface.
V MINERVA-II-1 was developed by the team thatdeveloped MINERVA, which was aboard thefirst Hayabusa. It comprises two rovers,
Rover-1A and Rover-1B.
V MINERVA-II-2, which carries Rover-2, is an
optional device developed by a university
consortium.
(© JAXA) 	

MINERVA-II-2	

Rover-2	

(Optional)	

MINERVA-II-1	

Rover-1A	

Rover-1B	',

MINERVA-II-1
V Successor to MINERVA, which was mounted aboard Hayabusa.
V Purpose: Engineering demonstration of the movement mechanism
V Development: MINERVA-II team (ISAS)
twith cooperation of the University of Aizu
V MINERVA-II-1 carries two (twin) rovers
V Mass including the deployment structure is 3.3 kg
tDimensions: 22.5 × 22.5 × 20.5 cm
V Rover mass: approx. 1.1 kg
tDimensions: diameter 18 × 7 cm
V Two cameras (wide-angle and stereo)
V Temperature sensor and photodiodeV Accelerometer and gyro
V Explorers move by hopping to explore the asteroid surface(© JAXA) 	

'-MINERVA-II-2
V Explorer robot developed by a university consortium
tThis is an optional, piggy-back device
V University consortium led by Tohoku University, co-developed with Tokyo
Denki University, Osaka University, Yamagata University, and the Tokyo
University of Science
V Total mass including separation mechanism: 1.6 kg
ttDimensions: 17.5 × 17.5 × 20.5 cm
V Rover mass: approx. 1 kg
ttDimensions: diameter 15 × 16 cm
V Mounted equipment include a camera, thermometer, photodiode, andaccelerometer
V Four types of mobility systems are equipped:
ttEnvironmentally dependent buckling mechanism (Yamagata University)
ttLeaf-spring buckling mechanism (Osaka University)
ttEccentric motor-type micro-hop mechanism (Tohoku University)
ttPermanent magnet-type impact generation mechanism (Tokyo Denki
University)
V The exploration robot hops to move across and explore the asteroid
surface.(© JAXA) 	

($

MASCOT
Mobile Asteroid Surface Scout 
V Created by DLR (German Aerospace
Center) and CNES (French National Centre
for Space Studies)
V Small lander with mass approx. 10 kg
V Carries four scientific instruments
V Can move only once, by jumping
Scientific instruments aboard MASCOT
DeviceFunction

Wide-angle camera(CAM)

Imaging at multiple
wavelengths

Spectroscopic microscope(MicrOmega)

Investigation ofmineral compositionand characteristics

Thermal radiometer(MARA)

Surface temperature
measurements

Magnetometer (MAG)

Magnetic ﬁeld
measurements

Flight model(© DLR) 	

MASCOT(%Electric propulsion (ion engine)
V Name: µ10
V Converts xenon* into plasma (ions), which isaccelerated by applying voltage.
V A microwave discharge system is used to generateions.
V Four units are mounted, and simultaneous operationof three generates thrusts of up to 28 mN.
V Approximately 60 kg of loaded xenon fuel, allowing
acceleration up to 2 km/s.

Injection test in a flight model
vacuum chamber

V It is used to alter trajectories when cruising from
Earth to the asteroid and back.

*Why we use xenon
V Xenon is a monoatomic molecule, so its ionization voltage
is smaller than that of gasses comprising two or more
atoms. This increases the ratio of added energy that isused for acceleration.
V Reactivity is lower than that of other substances.
V Mass (atomic weight) is large, improving the efficiency ofacceleration.

Hayabusa2 ion engine(© JAXA) 	

(&

Reference: How ion engines work
イオン生成

静電加速

中和

[[Translation on
next slide]]	

推進剤ne プラズマ
+ e
放電

n

+

e

e
e
+
+
+
e プラズマ

+

n
n

推進剤

n

e

n

+

イオン源

電位

+
+

+

+

+

+

e

+

+
e

+

+

アクセル
グリッド

e

e

噴射速度
30km/se++e

DC discharge method

+

e

e

ディセル
グリッド

+

e

+
+

スクリーン
グリッド
e +

e

+

+

放電

e

中和器

+
+

+
e

+

+

+

e

+

e

Source: Dynamic Navigation with Ion
Engines (Space Engineering Series 8),
Corona Publishing (2006)

Different ion generation systems

Cathode for thermionicemission
Note: The ion engine developed in the U.S., the U. K. and theformer NASDA was a DC discharge Kaufman-type ion
engine or a Ring-Cusped ion engine.

Microwave discharge method

Microwave generator
Note: The ion engine developed at the ISAS in Japan is a
microwave discharge-type ion engine.

('Ion production	

Electrostatic
acceleration	

Neutralization	

イオン生成

静電加速

中和

Propellant	

推進剤ne プラズマ
+ e
放電

n

+

e

e

+

e

+
+
Plasma	+e プラズマ
+

n
n

推進剤

n

e

n

Propellant	+

イオン源
Ion source	

電位Potential	

+

+

+

+

+

+

+

+

e

+
e

Neutralizer	

中和器

+
e

+

+

スクリーン
Screen grid	

グリッド
e +

+

e

e

アクセル ディセル
Acceleration
Deceleration
グリッド
グリッド
grid	

grid	

+
e

e

Thrust rate	

噴射速度30km/se++Discharge	

放電

e

e

+

+

Discharge	

e

+

+

+

Plasma	

+

e

e

+

e

+

((

Chemical propulsion systemV The chemical propulsion system is used for
attitude control (reaction wheel unloading,safe hold), fine trajectory modifications, andorbital control at the asteroid.
V The thruster is a 20 N two-component systemusing fuel (hydrazine) and an oxidizer(MON-3).
V There are 12 thrusters in total: 4 on the upper(+Z) surface, 4 on the lower (–Z) surface, 2 on
the surface with the ion engine (+X), and twoon the surface with the capsule (–X).
V The thruster system has a redundantconstruction.
V Approximately 48 kg of propellant is carried.

(© JAXA) 	

Red circles indicate thruster locations.
Not displayed are one thruster on thebottom surface, and one on each of
the opposing (ion engine) surfaces
between their upper and lower edges,
for a total of twelve thrusters.

()Chemical propulsion system: Changes from Hayabusa
■ Countermeasures against leaks that occurred immediately after Hayabusa touchdown (second time)_ Improved valve cleaning methods and airtightness tests, fewer welding locations, review of welding
procedures, etc.
■ Countermeasures against freezing in both pipe systems that occurred after Hayabusa leak
_ Separation of piping routes for the A and B systems and independent heat control■ Countermeasures against orbital insertion failure by the Akatsuki Venus orbiter_ Full separation of fuel and oxidizer pressure regulating systems■ Measures for realization of the Hayabusa2 impactor mission_ Confirmation of long-term thrust (collision avoidance) and short-pulse thrust (landing within craters)■ Other changes_ Metal diaphragm oxidant tank changed to a surface tension device**What is a surface tension device?
V This device uses helium gas to apply pressure when extracting
oxidant from its tank, ensuring that only oxidant fluid, not helium
gas, is extracted. Its naming comes from the fact that it utilizes
surface tension of the oxidant.

(© JAXA) 	

Test burn of a flight model to confirm longterm and short-pulse thrust.

(*

Attitude and orbital control system (AOCS¨V The AOCS is responsible for attitude control of the probe and navigation near theasteroid.
V Component devices are described below.
c Attitude and orbital control` Attitude detection sensor
 Coarse Sun Aspect Sensor (CSAS)
 Star Trackers (STT) Inertial Reference Unit (IRU) Accelerometer (ACM)

 Reaction Wheel (RW) Reaction Control System (RCS)d Other navigation equipmenta Asteroid relative position measurement sensor

 Laser altimeter (LIDAR) Laser Range Finder (LRF) Flashlight (FLA) Target Markers (TM)
 Drive (DRV)b Image processing component Optical Navigation Camera (ONC)
 Digital electronics (ONE-E)
AOCUª Attitude and orbit control unit
AOCPª Attitude and orbit control processor(+Target Markers
V Target markers descend to the satellite
surface before touchdown as artificial
landmarks. The explorer descends while
flashing a strobe to recognize the target
markers.
V Markers are fashioned like beanbags, with alarge number of pellets in a soft enclosure,
to prevent the marker from bouncing on theasteroid surface.

10 cm

V The outer material is highly reflective.
V Hayabusa2 carries five target markers(Hayabusa carried only three).
V Thin sheets with names inscribed arecontained within.

(© JAXA) 	

Shine whitein sunlight(,Electrical systemV In sunshine, electric power generated by the solar array paddles is supplied to onboard
equipment while the battery is charged. In the shade, the battery stably supplies
equipment with power throughout the mission.
V Following the Hayabusa design, this system provides reliability and improved power
supply. An outline of primary power supply system equipment is shown below.
■ Solar Array Paddles (SAP¨- Converts sunlight into electricity for supply to mounted
equipment- A high-efficiency 3-junction solar cell is used
- 3-panel × 2-wing construction produces 1460 W @1.42 AU■ Series-switching regulator (SSR¨- Stabilizes and controls SAP-generated power for supply tomounted equipment via the PCU■ Power control unit (PCU¨- Distributes and controls power from the SSR to mounted
equipment- Controls and manages power for recharging the BAT
■ Battery (BAT¨- Provides power through the PCU as needed while in shade,etc.
- Eleven inline-mounted 13.2 Ah lithium ion batteries

External view of the SAP(left: stored; right: deployed)

(© JAXA) 	

(-Deployment test for SAP deployment(© JAXA) 	

)$Communications (antennas)
V X-band (8 GHz) waves are
generally used forcommunication with groundstations.

°-band low-gain antenna(X-LGA-A¨
Ka-band high gain antenna(Ka-HGA¨

V There are three types of Xband antennas: high-, mid-,and low-gain.
V The Ka-band (32 GHz) is used
to transmit data from scientificobservations to Earth after
arrival at the asteroid.
V Approximately four times moredata can be transmitted via theKa-band than by the X-band.
However, transmissions arehighly affected by weather(attenuation due to rain ishigh).
V Bit rates are 8 bps–32 Kbps.

X-band high-gain
antenna(X-HGA¨°-band mid-gain antenna(X-MGA¨°-band low-gain antenna(X-LGA-B¨°-band low-gain antenna(X-LGA-C¨(© JAXA) 	

What about Ka-band reception?
Ka-band radio waves from planetary explorers currently cannot bereceived at tracking stations in Japan, so we use overseas tracking
stations.

)%Re-entry capsule At the very end of the Hayabusa 2 mission, a capsule carrying a container filled with asteroid samples will
re-enter the Earth’s atmosphere at 12 km/s and be collected on the ground.
 The capsule separates from the spaceship while spinning at one revolution per 3 seconds. It gets very hot
due to atmospheric entry (in technical terms, it passes through a corridor with aerodynamic heating of 14
MW/m2). It opens a parachute at an altitude of about 10 km, allowing it to gently descend and land while
outputting a beacon signal for positional search.

 Fundamental design is nearly the sameas that in the first Hayabusa, but
mounted equipment, parachutedeployment trigger (signal), andreliability of associated equipment havebeen improved.
 The Riparian Environment Management
Model (REMM) is newly added, and willmeasure acceleration, rotation, andinternal temperatures during flight.

(© JAXA) 	

Rear heat shield (integrated
with parachutecover)
AFT-TPS(Parachute
Cover)(Carbon Phenolic)Sampler
container~|
SampleContainer

Mass:
W=17
W=17
kg
approx.
16 kgkg
W: 約16.5kg
Dia=400mm
Dia=400mm
Diameter:Dia:H=200
400mm
mm
mm
approx.H=200
400 mmH=
200mm
Height: 
approx. 200 mm

Electronics
ParachuteParachute Insulator
FWD-TPS(CarbonPhenolic)
Front heatshield


 
Insulation
(Carbon Phenolic))&Re-entry capsule
Clothparachute (with laserPeT\bSahW
@3K31B7D31:=B6	iHdhRj.2`Zh[Avoidance maneuvers	

Separates from
T_f:(@%5F>-;
spacecraft
while
EGDL +KM
6
A>7<<7<5 @3D A	

Re-entry
approx. 12 km/s	

' speed:0l>?AB<C
Acquisition and ranging by

Parachute\bSahW4
i9(>=ABj
deployedkdhUHKNg!,primary radar	(approx. alt. 109;	Start4
re-entry'
/>>@=F9;/:B 	

i9(?==ABjAerodynamic&"heating	What is the temperature at re-entry?
V The capsule surface (thestagnation point of the heatresistant material) is at most 3000], but the capsule interior reaches
50 ] at most.

(© JAXA) 	

#8
Dynamic
instability
i170jregion	(range of transition
B=A=<71A>332A	]hWShcXHeatshield6
separation	

Beacon antenna extension^hQfOfVY3^hQf/
and signal transmission	

Landing (beacon continues$i$J^hQf/I)*j
B@/<A;7AA7=</4B3@:/<27<5	DesertRe-entry sequence overview)'Others
Detailed descriptions of the following systems are omitted here:
V Structural system: Overall spacecraft support
V Thermal control system: Manages spacecraft temperatures
V Data processing unit: Processing and control of all data
V Electric instrumentation: Wire-connecting equipment
V Digital Electronics (DE): Processes data from scientific sensors(ONC, TIR, NIRS3, DCAM3))(3. History of the mission))History (overview)
FY2011–2014
3 Dec 2014ª Development phaseª Launch

3–5 Dec 2014
6 Dec 2014–2 Mar 2015ª Critical operationsª Initial function check

Mar 2015–ª Cruising phase

3 Dec 2015
4 Dec 2015–Apr 2016ª Earth swing-byª Southern hemisphere station operations

22 Mar–21 May 2016
22 Nov 2016–26 Apr 2017ª phase-1 ion engine operationª phase-2 ion engine operation

10 Jan–3 Jun 2018ª phase-3 ion engine operation

27 Jun 2018

: Asteroid arrival)*LaunchV Rocketª H-IIA-26 (type 202)
V Planned launch date: 30 Nov 2014 (Sun) 13:24:48
ttttttttttttt←Delayed due to weather
V Actual launch date: 3 Dec 2014 (Wed) 13:22:04
V Possible launch windowª30 Nov–9 Dec 2014
V Launch location: Tanegashima Space Center
V Sub-payloads accompanying launch:
tttShin‘en 2 (Kyushu Institute of Technology)
tttARTSAT2-DESPATCH (Tama Art University)
tttPROCYON (co-research by University of Tokyo and JAXA¨)+H-IIA launch
Vehicle(/B3::7B34/7@7<5BG>3(	Satellite fairing	

Hayabusa 2	=>/G:=/2A	2-stage liquid-fuel rocket
Type H2A202

Stage 2 liquid hydrogen tank	Stage 2 liquid oxygen tank	Stage 2	Full length	

Stage 2 engine	Stage 1 liquid oxygen tank	Stage 1	Stage 1 liquid hydrogen tank	

Solid rocket booster	Solid fuel rocket
boosters	 </;7<5 /01/ AB <2AB/530<C;03@=4"'1<C;03@=4('	Stage 1 engine	(© JAXA) 	

),Rocket flight plan
Stage

Hours

Time after launch	

Minutes	

Seconds	

Altitude

Inertial
velocity

1. Liftoff	

2. Solid rocket booster burn completes*	

3. Solid rocket booster burn separates**	

4. Satellite fairing separation	

5. (B/53 3<57<30C@<AB=># 	

6. Stage 1/2 separation	

7. (B/53 >@7;/@G3<57<3AB/@B(  	8. (B/53 >@7;/@G3<57<3AB=>(	

9. (B/53 A31=<2/@G3<57<3AB/@B(  	10. (B/53 A31=<2/@G3<57<3AB=>(	

11. Hayabusa 2 separation	

12. Shin’en 2 separation	

13. ARTSAT2-DESPATCH separation
14. PROCYON separation	

*At burn chamber max. pressure 2%	

**Thrust strut cutoff	)-Rocket flight route&@7;/@G3<57<30C@<AB=>A# 	Geodetic latitude [northern latitude, deg]

Satellite fairing separation	

(B/53 >@7;/@G3<57<3AB=>(	Shin’en 2 separation	

PROCYON separation	

Hayabusa2
separation	ARTSAT2-DESPATCH separation	

(B/53 A31=<2/@G3<57<3AB/@B(  	

(B/53 A31=<2/@G3<57<3AB=>(	Geodetic longitude [eastern longitude, deg]*$Critical operations, initial function checkCritical operations (3–5 Dec 2014)
V Solar array panel deployment, sun acquisition control
V Sampling device horn extension
V Release launch lock on the retaining mechanism for the gimbal that controls
ion engine direction
V Confirm spacecraft tri-axial attitude control functions
V Ground-based confirmation of functions for precise trajectory determination
systemInitial functional confirmation (6 Dec 2014–2 Mar 2015)
V Confirmation of ion engine, communications, power supply, attitude control,
observation devices, etc.
V Precise trajectory determination*%Initial function check (details)
Date

2014

Tasks performed

12/7,8Functional confirmation of X-band mid-gain antenna beam pattern measurements, acquisition of actual data, and X-band communicationequipment

12/9

Power system (battery) function check

12/10

Near-infrared spectrometer (NIRS3) inspection

12/11

Inspection of thermal infrared camera (TIR), deployable camera (DCAM3), Optical Navigation Camera (ONC)

12/12–15

Function check for attitude and trajectory system (all devices)

12/16

Inspection of miniature rover (MINERVA-II) and lander (MASOT)

12/17

Inspection of re-entry capsule and impactor (SCI)

12/18

5-point pointing test of X-band high-gain antenna (XHGA), pre-operation of ion engine

12/19–22

Ion engine baking

12/23–26

Ion engine test operation (ignition) *performed for each engine[12/23: ion engine A; 12/24 ion engine B; 12/25: ion engine C; 12/26: ion engine D]

12/27–1/4

Precise trajectory determination, Delta Differential One-way Ranging (DDOR)[No operations on 12/28, 1/1–2]

2015

1/5–7

Ka-band communications device actual data acquisition, antenna pattern measurements

1/9–10Ka-band DSN station DOR, lensing tests

1/11

Ion engine pre-operations

1/12–15
1/16
1/19–20
1/23

1/20–3/2

Ion engine paired test operations [1/12: A＋C; 1/13: C＋D; 1/14: A＋D; 1/15: A＋C]
Ion engine tri-set testing: A＋C＋D
Paired engine 24-hour continuous autonomous operation: A＋D
Function check of laser altimeter (LIDAR), laser range-finder (LRF), flash lamp (FLA)
Confirming functions such as coordinated operation of multiple devices for transition to cruising phase (regular operations)
Function check of linked operations, such as solar light pressure effects evaluation, data acquisition from sun tracking movement behavior,solar light pressure and attitude trajectory control equipment (reaction wheels, etc.), ion engine*&Mar 2015 swingby
2015.03.02

Initial operations phase complete, followed by normal
operations phase.

2015.03.03–21
2015.03.27–05.07

EDVEGA phase-1 IES operation
Solar sail mode operations(maintains fuel-free solar orientation using only 1 RW out of 4.
Other RWs are kept in the OFF state)2015.05.12–13

Three IES operate in 24-hour mode (ITR-A+C+D¨

2015.06.02–06
2015.09.06

EDVEGA phase-2 IES operation
Solar sail mode operation starts

2015.09.01–02

IES-TCM (precise trajectory control for swing-by)

2015.10.01–12.03
2015.12.03

Precise guidance phase (TCM by RCS twice)
Earth swing-by*'After Earth swingby through late 2016
–2016.04E
2016.03.22
2016.05.21
2016.05.24, 06.01–09
2016.06.14–20
2016.06.22–23
2016.06.29–07.03
2016.07.05–08
2016.08.03
2016.10.08
2016.10.11–16
2016.10.19–22
2016.11.02, 04
2016.11.22

Southern hemisphere station operations (by DSN Canberra
and ESA Malargüe only)
Transfer phase-1 ion engine operations startTransfer phase-1 ion engine operations end
Mars observations (–Z Mars orientation)
Light pressure confirmation operations
DSN–DSN uplink transfer testing
DSN Ka-band communication testing
ESA Ka-band compatibility testing
Transition to attitude control solar sail mode
Transition to 3-axis attitude control wheel
STT Mars observations (OPNAV practice)
ONC fixed-star observations
DSN–UDSC uplink transfer testing
Transfer phase-2 ion engine operations start*(2017–2017.04.18
2017.04.26
2017.05.16–28
2017.05/30–06.01
2017.09.05
2017.11.18, 28
2017.12.02
2017.12.26–27
2018.01.10
2018.02.26
2018.06.03
2018.06.03
2018.06.27
ttª

ONC-T imaging near L5
Transfer phase-2 ion engine operations end
ONC imaging of Jupiter and fixed stars
RCS autonomous maneuvering tests
Reset internal clock (TI) to zeroDSN–SSOC real-time Doppler transmission testing
DSN–UDSC uplink transfer testing
IES test maneuvers
Transfer phase-3 ion engine operations startFirst Ryugu observations
Transfer phase-3 ion engine operations end
Asteroid approach navigation start
Asteroid arrival*)Summary of ion engine operationsPhase-3 ion engine operations(2018.01.10–06.03)
Trajectory to Ryugu
Hayabusa 2 orbit

Phase-2 ion engine operations(2016.11.22–2017.04.26¨SunEarth orbit■ Before swing-by
Period

Name

Units

Accel.
m/s

Time
H

Initial functioning
confirmation

IES operations testing

-

−

−

2015.03.03–21

IES powered flight 1

2

44

409

2015.05.12–13

IES max. thrust test

3

4

24

2015.06.02–06

IES powered flight 2

2

11

102

2015.09.01–2

IES powered flight 3

2

1. 3

12

Launch(2014.12.03)

Arrive at Ryugu(2018.06.27)

IES：Ion Engine System	

Phase-1 ion engine operations(2016.03.22–05.21, incl. added burns)

Earth swingby(2015.12.03)

■ After swing-by(© JAXA) 	

Period

Name

Units

Accel.
m/s

Timeh2016.03.22–2016.05.21

Phase-1 ion engine operations

3 (2 at times）

127

798

2016.11.22–2017.04.26

Phase-2 ion engine operations

3 (2 at times）

435

2593

2018.01.10–2018.06.05

Phase-3 ion engine operations

2→3

393

2475

**

Description of primary operations*+Solar sail mode (2015–)
Attitude control using the power of sunlight
A new technology that requires only a single reaction wheel; no fuel needed
V A new technology for Hayabusa2 that utilizes findings from Hayabusa and IKAROS
V This technology (a type of “solar sail” technology for utilizing the power of sunlight) allows
stable control of spacecraft attitude with only one of the four reaction wheels aboard
Hayabusa 2 turned ON, others OFF.
V Realizes non-fueled, long-term maintenance of sunward orientation, which was not possible
in earlier spacecraft.
^Attitude maintenance realized by this technology for over 9 months of the 2.5-year flight.

RW-X

Only one reaction wheel(RW-Z1) left ON

(© JAXA) 	

RW-Z1

RW-Y

RW-Z2

Hayabusa(2003–2010¨

IKAROS(2010–¨*,Scientific results from the swing-by (3 Dec 2015)
ONC-T

TIR

Color Earth image

Intensity distribution oflight reflected fromvegetation

NIRS3

Australia (warmer than ocean)

TIR thermal image

ONC-T color image

LIDAR
Moonstrongwave length(µm)

signal level 

Earth

data NO

Light absorption by water molecules inEarth's atmosphere(© JAXA) 	

weak

Successful laser reception at 6.7 million km (0.045 AU)
on 19 Dec 2015*-Mars imaging (May–Jun 2016)
 24 May, 1–9 Jun 2016
 We performed observations, taking advantage of an alignment of Hayabusa2, Earth, andMars. (Observations by ONC-T, NIRS3, TIR)

Near alignment	

Venus	

Mercury	7AB/<13*	

Earth	

Hayabusa 2	

Mars	

ONC-T image of Mars
21:46 24 May 2016 (Japan time)

Ryugu	

7AB/<13*	(© JAXA) 	

70

Uplink transfer (Jun–Nov 2016)
Uplink transfer technology testingª22–23 Jun 2016 ← between DSN stations
ttttttttttttttttt 2–4 Nov 2016t← between Usuda–DSN
First successful
test in Japan!

Previous method

Station A

Communications
temporarily cut

Station B

Uplink transferªCommunications
not cut(© JAXA) 	

Station A

Station A

Communications
not cut
Station B

Station B

+%

	

Ka-band communications, DDOR (Jun–Jul 2016)
Ka-band technology testing: 29 Jun–8 Jul 2016

V 29 Jun–3 Jul 2016: Ka-band communications testing at DSN Stn (Goldstone) t← success from approx. 50 million km!V 1–2 Jul 2016: Ka-band DDOR testing between NASA–ESA stations (NASADSN: Goldstone, ESA: Malargüe)
tt←World-first Ka-band DDOR between 3 organizations!

V 5–8 Jul 2016: Ka-band communications testing at ESA station°-band low-gain antenna(X-LGA-A¨

X-band (8 GHz): Normal operations

Ka-band high gain antenna(Ka-HGA¨

Ka-band (32 GHz): Can transmit approx. 4times more data than X-band. Used tosend asteroid observation data to Earth.

X-band high-gain
antenna(X-HGA¨°-band mid-gain antenna(X-MGA¨

Ka-band is rarely used in deepspace exploration

°-band low-gain antenna(X-LGA-B¨°-band low-gain antenna(X-LGA-C¨+&DDOR

	

DDORªDelta Differential One-way Ranging
At least two ground stations simultaneously receive radio waves from the spacecraft. In
addition, we receive radio waves emitted from a visible celestial body (a quasar) that is as
visually close as possible to the spacecraft. By comparing data received at two or moreground stations, the probe trajectory can be determined with high accuracy. (Radio waves
from the probe and those from the quasar are received alternately.) This is the same
principle as VLBI.

QSO

Usuda	

Canberra	

*Blue arrows are signals from a quasarGoldstone	

By simultaneously acquiring data along east–west and north–south baselines, we
succeeded in determining high-precision
trajectories during ion engine operations(micro-thrust accelerations)! A world's first!

+'Imaging at L5 (18 Apr 2017)
Observation
 Date: 18 Apr 2017 (Japan time) Three sets of four continuous images at 30 min intervals from the Optical
Navigation Camera (ONC−T) telescope Exposure time: 178 sec (longest exposure)
Results
 No moving objects were seen in any sets

Orbital direction	

Sun	

Earth swingby	

Asteroid
arrival	

Launch	

Earth	

Loc. At	

17 Mar 2017	

Sun–Earth L5	

Sun–Earth system Lagrange points L4, L5(© JAXA) 	

+(

Jupiter observation (16–17 May 2017)

 Date: 16 May 2017 17:30 (universal time)

tttttttt17 May 2017 02:30 (Japan time) View angle:tt0.79 × 0.79 deg
 Exposure time: 0.1312 s
 Wavelength:ttt v-band (550 nm)
 Distance to Jupiter (16 May 2017 17:30 UT):

ttttttttt4.48565 aut(6.71044 x 108 km)

 Magnitude as seen from spacecraft: –2.44
 Imaging objective:\

Jupiter as imaged by ONC-T

Various devices aboard Hayabusa2 perform observations in
preparation for arrival at the asteroid about one year later. The
figure shows a calibration observation for the visible
spectroscopic camera, targeting Jupiter as the brightest planet.

(© JAXA) 	

+)

TI reset (5 Sep 2017)
 Time (TI) reset of the spacecraft clock Clock is reset through operations on 5 Sep 2017
 No need for further resets until return to Earth■ Description
V Spacecraft-internal time counter: 32 bits
V Time count: 1 count « approx. 31 ms (1 ms «1/1000 s)
V 32 bits allows counting to 4,294,967,296 (approx. 4
yr 3 mo)
V Counter reverts to zero after reaching max value(like a car odometer)
V This is performed to avoid a counter value of zero
during stay at Ryugu

Value before counter reset	Value after counter reset	+*First observation of Ryugu (26 Feb 2018)

Successful imaging of Ryugu by

the onboard ONC-T camera on 26
Feb 2018

Ryugu	

Observation conditions were goodon this day; Ryugu was in the
ONC-T FoV without making large
attitude corrections.

Distance from spacecraft to Ryugu
was approx. 1.3 million km

Three images are overlaid. Ryugu ismoving in the direction of the pink arrow.
View angle in the image is 0.8 deg)

(ONC team: JAXA, Univ. Tokyo, Kochi Univ., Rikkyo Univ., Nagoya Univ.,
Chiba Inst. of Tech., Meiji Univ., Aizu Univ., AIST)

++

4. Trajectories+,Trajectories overviewAfter launch, the spacecraft enters a trajectory close to Earth orbit, and returns to Earth for aswing-by exactly 1 year later. After the swing-by, it enters a trajectory close to orbit of asteroid
Ryugu, arriving there after about two orbits. It will remain at Ryugu over a little more than one
revolution around the sun. After that, it will leave Ryugu, revolve around the sun for a little morethan one orbit, then return to Earth.
Hayabusa 2 trajectory
Ryugu orbit

Event

Date

Launch3 Dec 2014

Earth swing-by

3 Dec 2015

Asteroid arrival

27 Jul 2018

Asteroid departure

Nov - Dec 2019

Return to EarthNov - Dec 2020

Earth orbit
Sun
Arrival at Ryugu(27 Jul 2018)(© JAXA) 	

Earth swing-by(3 Dec 2015)

Launch(3 Dec 2014)

Earth departure to asteroid arrival+-Trajectories: Launch to Earth swing-by```: Dec 2014
aaa : Feb 2015``abbb : Apr 2015
ccc : Jul 2015dea

ddd : Sep 2015ee

eee : Nov 2015
b

d

b

d c

Red: Hayabusa2

c
`

c
a

(© JAXA) 	

b

Blue: Ryugu
Green: Earth

Launch near `, return and Earth swing-by near e. There is little
distance between Earth and Hayabusa2.

,$Trajectories: Earth swing-by to first orbit

fff: Dec 2015
ggg: Feb 2016ff

hhh: Apr 2016

kiii: Jul 2016

f
gg
j
g

h
h

i

j

i
h

(© JAXA) 	

k

jjj: Sep 2016kkkk: Nov 2016j

Red: Hayabusa2
Blue: Ryugui

Green: Earth

After Earth swing-by near f, Hayabusa2 leaves Earth and gradually
approaches Ryugu (at k).

,%Trajectories: First to second orbit (asteroid arrival)

lll: Jan 2017
mmm: Apr 2017
⑭

⑬

⑬
⑱

⑰

⑭

⑮

nnn: Jun 2017

⑬

⑲
⑭

⑯

⑮

⑲⑲

ooo: Aug 2017
ppp: Nov 2017
qqq: Jan 2018⑱⑱

rrr: Mar 2018

⑮
⑯

⑯

⑰⑰

Red: Hayabusa2
Blue: Ryugu
Green: Earth(© JAXA) 	

While making one more orbit from l to r, Hayabusa2 makes one
more orbit while approaching Ryugu.

,&Trajectories: Stay at asteroid
sss: Jun 2018
: Aug 2018
: Oct 2018

⑳⑳
㉗㉗

㉑㉑

: Jan 2019

㉘

㉖㉖

㉓

㉒

㉔

: Mar 2019㉗

: Jul 2019

㉘㉘

㉕㉕
㉑
㉒㉒

: May 2019㉕⑳

: Oct 2019¡¡¡: Dec 2019

㉖
㉔㉔
㉓㉓

Red: Hayabusa2
Blue: Ryugu
Green: Earth(© JAXA) 	

Hayabusa2 arrives at Ryugu near s, and travels with the asteroidfor over one solar orbit to ¡.

,'Trajectories: Asteroid to Earth¡¡¡: Dec 2019¢¢¢: Feb 2020£££: Apr 2020㉘¤¤¤: Jul 2020㉙¥¥¥: Sep 2020㉘㉘¦¦¦: Nov 2020
㉚
㉛
㉙㉙

㉛㉛
㉚㉚

(© JAXA) 	

Red: Hayabusa2
Blue: Ryugu
Green: Earth

Hayabusa2 departs Ryugu at around ¡, then heads directly to Earth toreturn the capsule near ¦.

,(Trajectories in rotational coordinates 

Earth departure
Earth swingby
Ryugu arrival
Ryugu departure
Earth re-entry

3 Dec 2014
3 Dec 2015
27 Jun 2018
Nov–Dec 2019Nov–Dec 2020

C3 = 21 km2/s2
Ion engine total impulset
ttttt= 2 km/sec
Re-entry speed = 11.6 km/s

Ryugu	
  departure	
  Nov–Dec	
  2019	Trajectory	
  for	  Earth	
  return	
  from	  Ryugu	

Ryugu	
  orbit	

Earth	
  re-­‐entry	  Nov–Dec	
  2020	

Opera�onal	
  trajectory	
  in	
  Ryugu	
  vicinity	

Earth	
  swingby	
  Dec	
  2015	

Total flight time = 6 yr (4.5 yr cruisingtime)
Total powered flight time = 1.5 yr
Total flight distance « 5,240,000,000
km

Arrival	
  at	
  Ryugu	
  Jun–Jul	  2018	EDVEGA	  loop	

Transi�onary	
  trajectory	
  from	  Earth	  to	
  Ryugu	(1999 JU3 = Ryugu¨(© JAXA) 	

,)

Earth swing-by
V Hayabusa2 approached Earth for a swing-by on 3 Dec 2015.
V Earth approach time: 19:08 (Japan time)
V Passed approximately 3,090 km over the Hawaiian islands
Hayabusa2 trajectory

Overview of orbits

Ryugu orbit

Swing-by increases flight
speed by 1.6 km/s.
Solar relational speed attime of swing-by changed
from 30.3 km/s to 31.9
km/s.

Sun
Arrival at Ryugu(27 Jun 2018)

Earth swing-by(3 Dec 2015)

Earth orbit
Launch(3 Dec 2014)

(© JAXA) 	

86

Swing-by trajectorySolar system view from north
Diagrams depicting orbits around the sun. These figures show orbits of Earth and Hayabusa2
around the sun. The degree of curvature of the Hayabusa 2 orbit at the swing-by point thus
appears small.
Swing-by point

Earth

Hayabusa2 Hayabusa2 (© JAXA) 	

Before swing-by (Sep 2015)

Ryugu

After swing-by (Jun 2016)

87

Operations before and after swingby	

2 Nov 2015 TCM1
10–13 Nov 2015
Earth and lunar imaging by thermal
infrared camera

26 Nov 2015
TCM21 Dec 2015
TCM3―Cancel
3 Dec 2015
Closest Earth approach(swingby)
Lunar orbit

26 Nov 2015
Earth and lunar imaging by opticalnavigation camera (telescope), thermal
infrared cameraunear-infrared
spectrometer
3 Dec 2015
Earth imaging by optical navigation
camera (wide-angle),
imaging tracking function checkNorthern polar
direction
Enter shade(18:58 JST)

Sunshade (20min)

Sunward direction
Closest approach
4 Dec 2015
(19:08:07 JST)
Earth imaging by optical navigation
camera (telescope)and thermal infrared camera

Sunward direction
Leave shade(19:18 JST)

19 Dec 2015
LIDAR optical link experiment

22 Dec 2015
Cancel Earth observation attitude
and transition to cruise attitude

Primary operations before and after Earth swing-by(© JAXA) 	

(Times are JST)

Trajectories at closest Earthapproach88Principle of swing-bys
Earth region escape velocity
Solar escape velocity

¯²

¯­
Change in velocity relative tosun
Velocity ¯¬ becomes VE

Velocity of Earth rotation

¯­

¯¬

¯±
Approach velocity to Earth region(Earth at center)
Approach velocity with respect to sun¯¬(© JAXA) 	

89

Simple description of swing-bysBall
10 m/s

14 m/s
14 m/sCar

10 m/sThrow a ball at 10 m/s at aright angle toward a car
travelling at 10 m/s.

From the perspective of thedriver, the ball approaches
the car diagonally atapprox. 14 m/s.

The direction of ball travel has changed by 90 deg,and its speed increased from 10 m/s to 24 m/s.

In this metaphor, Hayabusa2 is the ball, and Earth andits gravity are the car and the driver

The driver catches the ball,and throws it at 14 m/s in the
direction of travel.

14 m/s

10 m/sThe ball is now
travelling at 24 m/s withrespect to the ground90

5. Near-asteroid operations-%Sampling operation sequence`Separate from home position,
start GCP-NAV (combined
ground/onboard navigation)aEnter automatic modebDeploy target markers
cAttitude adjustment withregard to asteroid surface
dTouchdown
eShelter ΔV

Automatic/autonomous
technologies
tt||GSP, GCP-NAV

(© JAXA) 	

u

Ryug

-&

Automatic/autonomous technologiesªGSP, GCP-NAV
V Ground Control Point Navigation (GCP-NAV)ü Used for remote operations during approach from 20 km
to several hundred meters.
ü Satellite images transmitted to ground. By matching
feature points and contours of the asteroid with computer
generated template images, we can detect position andattitude information of the spacecraft and the asteroid.
ü Based on this, calculate levels of engine thrust on the
ground and issue commands to the spacecraft.
ü Human beings are good at recognizing complex images
and instantaneous judgments of the overall situation.
Ground instructions are thus advantageous despite thecommunication time lag.

V Guidance Sequence Program (GSP)ü From sensor information, autonomous behavior patterns performed by the spacecraft canbe efficiently rewritten and delivered from the ground.
ü We first obtain asteroid information that can only be derived through proximal observations,such as its surface conditions and reflectivity. Operators on the ground analyze thisinformation to determine risk assessments and how to handle emergency situations. Before
starting autonomous operations, ground commands are sent to rewrite tables in thespacecraft.
ü Efficient rewriting and instruction mechanisms are important for accommodating spacecraft
restrictions on communications capacity and computer memory.

Example of GCP-NAV operation screen(© JAXA) 	

-'

Impactor operations sequence` SCI separationa Horizontal
avoidance

b Vertical avoidanceImpactobservation

Lands here

c DCAM3
separation

Explosionand impact

Ryugu



Pinpoint
touchdownd Explosion andimpacte Return to home
position

(© JAXA) 	

-(

Impactor: Debris and ejecta avoidanceImpactor operates from above theasteroid (alt. several hundred meters)

Impactor separation
Explosion1) Debris avoidanceDebris rising due to explosion of theonboard impactor are avoided behind
the asteroid.

(a)
(b)

2) High-speed ejecta avoidanceAlso avoid high-speed ejecta produced
by projectile impact in 1).Ryugu	(© JAXA) 	

(a) High-speed debris(b) High-speed ejecta(c) Low-speed ejecta(c)3) Low-speed ejecta avoidanceLow-speed ejecta falling back onto
orbit around the asteroid are avoided
by maintaining sufficient distance. Lowspeed ejecta attaining very highaltitudes are sufficiently slow as to nothave a large effect, and chances ofcollision are low.

-)Pinpoint touchdownV Target Markers (TM)ü TM separate at an altitude of several tens ofmeters, and flash lamps intermittently illuminate
TM while cameras image them.
ü By comparing differences in images when flash
lamps are lit and when they are not, we canaccurately extract TM without effects from surface
patterns or sunlight.
ü Facing toward identified TM, descend to theasteroid while using laser altimeter information todetermine attitude and distance to the surface.
ü 6-degree-of-freedom (position + attitude) gas jet
injection control with high target tracking while
minimizing fuel consumption is also a key
technology.

Using one TM	

V Use of multiple TM

Using multiple TM	ü We will touch down near the artificial crater, andattempt to retrieve samples from exposed areas.

Crater	Range of error
with no other TM
for use as clues	(© JAXA) 	

Range of error fordescent to TM #1	

Range of error fordescent to TM #2
with TM #1 visible	

First
descent	

Seconddescent	

Range of error fordescent to TM #3
with TM #2 visible	

Third
descent	ü We expect the artificial crater to have a diameterof around several meters. By approaching thedestination point based on clues from multiple
sequential TM, we can perform the touchdownwith higher precision (a pinpoint touchdown).

-*

Spacecraft trajectory calculation near the asteroidConsidering the forces
described here, calculate thetrajectory of the spacecraft.

Earth

Sun
GravityKnown
Gravity

Solar radiation pressure*Because planetary
orbits around the sun
are well known

GravityPropulsion

PlanetEstimated

Gravity
Ryugu
(© JAXA) 	

*If gravity on Ryugu can be
estimated, we can learn its mass.
→ Density can be calculated whenvolume is known by shape estimation-+6. Operations-,Tracking stations used in operations(Illustration: Akihiro Ikeshita)

JAXA Sagamihara Campus
Uplink: CMD, RNG (7-GHz band)n Telemetry reception, confirmation, analysisn Command transmissionn Trajectory determination and predictionn Trajectory planning

Downlink: TLM, RNG (8-GHz band¨
Downlink: TLM, RNG (32-GHz band¨

Tracking
stations
Usuda 64 m
X-band(© JAXA) 	

Ka-bandMain stn

Uchinoura 34 m
Sub stn

Uchinoura 20 m
Launch support

ESA/DLRDSN 34 m, 70 m
NASA tracking support

DLR tracking support

Main stn--Tracking station locationsGoldstoneMalargüe

DLR, Weilheim
(Cebreros)
MadridUsuda
Uchinoura

Canberra(New Norcia)

( ): Planned usageUsuda is used for normal operations (with Uchinoura also used during launch)Critical operations also performed by the NASA Deep Space Network (DSN) (DSN: Goldstone, Madrid, Canberra)We are also trying to arrange use of the Weilheim tracking station with support of the German Aerospace Center(DLR) and the European Space Agency Tracking Station Network (ESTRACK) (Malargüe presumed).

%$$Tracking station locations (Part 2)
World map with Japan at center                    East longitude	



340

0

20

40

80

100

120

140

160

¨

180

200

220

240

260

280

300

320

340

0







(Cebreros) Goldstone
(ESA¨● DLR, Weilheim








(NASA¨  
Usuda 64 m


(JAXA¨ MadridUchinoura 34 m
 (NASA¨(JAXA¨(New Norcia)  Canberra(ESA¨

Malargüe  
(NASA¨(ESA¨


 




                   
80

80

70

70

60

60

 ( ESA ¨



50

50

 ( NASA ¨}  (®¬°¬¨ ( NASA ¨



40

40

30

30

(¨

 



North latitude	

60

(

20

20

10

10

- 10

- 10

- 20

- 20

ESA ¨

w(

- 30
- 40

~}(

{(

- 30

¨

- 40

NASA ¨

- 50

- 50

- 60

- 60

340

0

20

40

60

80

100

120

140

160

180

200

220

240

260

280

300

320

340

( ): Planned usage

0

%$%

Trajectory planning and determination
Spacecraft trajectory operations are conducted under cooperation between the orbit
planning group and the orbit determination group.

Trajectory planning

Trajectory determination

Before launch: Plan spacecraft trajectories fromlaunch to Earth return (consideringvarious conditions such as rockets,
ion engines, communications, andheat).

Before launch:
Confirm precision of trajectory
determination from launch to Earthreturn. Perform analysis presuming
data error.

After launch: Trajectory plans are correctedbased on orbit determination values,in particular adjustments to ion
engine operations.
 Cooperation with the attitude
systems group for trajectories near
the asteroid.

After launch: Based on actually acquired data*,
estimate spacecraft trajectory(positioning and velocity).
 Attitude systems group supports
trajectory estimations for trajectories
near the asteroid.

%$&Data used for orbital determination (1/2)
■ Normally, range and range rate (Doppler) data are used forspacecraft trajectory determinations.
Range rate (Doppler)

Range

1-way

We can know the distance to thespacecraft by sending radio
waves from a ground station andmeasuring the return time of radio
waves sent back. This is called
the “range.”

2-way

3-way

Both the spacecraft and the ground station are moving, so thefrequencies of radio waves traveling between them change due to the
Doppler effect, like any other wave. In other words, by examiningchanges in transmitted and received radio wave frequencies, the line-ofsight speed of spacecraft with respect to the ground station can be
known. This is called the “range rate” or “Doppler.” Distances can bemeasured using methods called 1-way, 2-way, or 3-way Doppler.

“Radio navigation” refers to estimating position and velocity (trajectory) of
the spacecraft using the range and range rate.

%$'Data used for orbital determination (2/2)
■ DDORª A technique called Delta Differential One-way Ranging is used to more accurately
determine trajectories.
At least two ground stations simultaneously receive radio waves fromthe spacecraft. In addition, we receive radio waves emitted from avisible celestial body (a quasar) that is visually as close as possible
to the spacecraft. By interfering data received at two or more groundstations, the probe trajectory can be determined with high accuracy.
(Radio waves from the probe and those from the quasar are received
alternately.)

Blue arrows show waves from the quasar

■Optical navigation: Optical navigation, in which data from spacecraft cameras supplement
radio navigation, is performed immediately before arrival at the asteroid.

Radio navigation
Asteroid
imaged bycameras

By imaging the asteroid using onboard cameras, we candetermine the direction to the asteroid as seen from thespacecraft. The spacecraft position can be accurately
determined through DDOR, but uncertainty regarding theasteroid’s position remains. Accurately approaching theasteroid while verifying its position from the spacecraft iscalled optical navigation.

%$(

7. Target body%$)Asteroid Ryugu

0TUGSQKF=ZVIVKOCIGFXKUJUJG;:2!?"?JGRJQUQISCRJXCTUCMGPQP7VPG&* &$%,CU
CSQVPF%&.)$7>?"?JGFKTUCPEGUQ=ZVIVKTCDQVU&&MO"6OCIGESGFKU.70B0 @PKWGSTKUZQH?QMZQ 8QEJK@PKWGSTKUZ =KMMZQ@PKWGSTKUZ :CIQZC@PKWGSTKUZ 2JKDC6PTUKUVUGQH?GEJPQNQIZ 9GKLK@PKWGSTKUZ @PKWGSTKUZQH0K[V 06>?"

%$*

Asteroid Ryugu
ttOrbit of Ryugu
:QUKEGª?JGFCUCTJQXPJGSGFQPQUKPENVFGUJGFCUCQDUCKPGFDZUJGTRCEGESCHU"
Name: Ryugu
Permanent designation : 162173
Provisional designation : 1999 JU3Discovered: May 1999
Sizet

: Approx. 900 m

Shape
Rotation periodt: Nearly spherical: approx. 7 h 38 min

Rotation orientation : Ecliptic longitude λ= 310°–340°
Ecliptic latitude β= –40°±–15°
Reflectivity
Typet
Orbital radius: 0.05 (blackish): Type C (assumed to comprise materials
containing water and organics): Approx. 180,000,000 km

Revolution cycle: Approx. 1.3 yr

Density and mass

: Density is currently unknown, but presumed
to be 0.5–4.0 g/cm3: Mass is approx. 1.7 × 1011 kg –1.4×1012 kg.

Estimated shape

(by T. Mueller¨%$+Asteroid Ryugu (detailed information)
:QUKEGª?JGFCUCTJQXPJGSGFQPQUKPENVFGUJGFCUCQDUCKPGFDZUJGTRCEGESCHU"
162173 Ryugu (1999 JU3) near-Earth asteroid (Apollo group)
Orbital elements: epoch 2458000.5 TDB (4 Sep 2017 0:00 UTC¨JPL Small-Body Database Browser
https://ssd.jpl.nasa.gov/sbdb.cgi#top, accessed 10 Dec 2017
 Semi-major axis 1.18956 au; eccentricity: 0.19028; inclination: 5.8839°
 Ascending node long.: 251.591°;targument of perihelion: 211. 447°;tperihelion passage: 13 Feb 2017.
25148
 Period: 473.8908 days « 1.29747 yr
 Perihelion distance: 0.96321 au; aphelion distance: 1.41592 au
 Minimum orbit intersection distance: 0.00112 au (potentially dangerous asteroid)
Physical parameters Rotation period: 7.6326 h; ecliptic longitude (λ¨325±15° ecliptic latitude (β¨-40±15° Thermal inertia: 150―300 J m-2 s-1/2 K-1, extremely low surface roughness [Müller+ 2017]
 Mean geometric radius: 865 ± 15 m, nearly spherical [Müller+ 2017]
 Albedo: geometric 0.047±0.003, Bond 0.014±0.002 [Ishiguro+ 2014]
 Spectral type: Cg [Binzel+ 2001]. The reflection spectrum gradient is nearly flat, but slightly reddened in
the near-infrared region and with a slight drop in the ultraviolet region. This resembles the reflection
spectra of CM and CI meteorites that have experienced heating. [Perna+ 2017]
(TDB: In solar system dynamics, 1 au = 1.49598×1011 m¨%$,Size comparison between Ryugu and Itokawa

Ryugu

Approx. 900 m(© JAXA) 	

Tokyo Skytree634m

Itokawa
535m

Tokyo Tower
333m

%$-

(162173) 1999 JU3 (Ryugu) spectrum

Relative magnitude

Relative reflectance

Belongs to another low-albedo family	

1999 JU3= Ryugu	The largest member of Erigone Family	

Wavelength [µm]
(Data from Viras (2008), Sugita et al. (2012), Abe et al. (2008))
%%$

(162173) 1999 JU3 (Ryugu) light curve
1.8

Model
Tp=7.625 hr assumed

Differential Magnitude

1.7
1.6
1.5
1.4
1.3
1.2

0

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9
Rotational phase1

(from Kim, Choi, Moon et al. A&A 550, L11 (2013))
%%%

History of Ryugu naming10 May 1999: U.S. LINEAR team discovers asteroid 1999 JU3 (provisional name) at

the Socorro observatory.
Oct 2006: 1999 JU3 is listed as a candidate exploration target in the “Hayabusa
Successor Spacecraft” proposal.
Aug 2013: Request for naming 1999 JU3 in the Hayabusa 2 project submitted to the
LINEAR team, and approval received.
22 Jul–31 Aug 2015: A campaign for collecting proposed names is conducted.
Approximately 7,300 suggestions are received, from which “Ryugu” is selected.
Sep 2015: LINEAR team submits the name “Ryugu” to the International Astronomical
Union.
28 Sep 2015: Ryugu is published in the Minor Planet Circulars as (162173) Ryugu =1999 JU3.

%%&Selection of asteroids for exploration
Conditions for selecting an asteroid for exploration:
 Scientific objectives
ttIn the Hayabusa 2 project, a C-type asteroid
Engineering requirements
ttAbility to return within Hayabusa 2 capabilities
tttt→ Limits on trajectory size and inclination
tt Ability for touchdown within Hayabusa 2 capabilities
tttt→ Limits on asteroid size and revolution period

Explorable asteroids
Orbit between those of Earth andMars, low inclination from Earthorbit (From spacecraft trajectory
control capabilities)
Revolution period of at least about
6 hours (From navigational
capabilities at touchdown)Diameter of at least several
hundred meters (To allow for
crater creation by impactor)

Note: We also searched for backup targets, but only 1999 JU3/Ryugu was found tobe appropriate.
%%'Required acceleration from Earth to asteroid and asteroid
absolute magnitude(© Y. Tsuda et al. Acta Astronautica 91 (2013) 356–362) 	

%%(Reference information

%%)Categorization by asteroid spectral type

Adapted by Usui from Bus & Binzel (2002)%%*Classification and ratios of asteroid types
Classification of asteroids by reflective spectrum

D-typeMore primitive
Tagish Lake meteorite?

Broadly S-type

S-type· Rocky· Ordinary chondrites(by T. Hiroi)

Broadly C-typeDistance from sun· Rich in organics and hydrated minerals· Carbonaceous chondrites

C-type(by D. Tholen in ASTEROID II)%%+Features of each type (1/2)
Type

Spectral form

Distribution

Correlation with meteorites

C

Flat from 0.45–0.9 µm.
Absorption of hydrous mineral origins in
3-µm band in most cases.

Often outside of the asteroid belt.

Carbonaceous chondrites

S

Reflectance increases from 0.4 to 0.7µm, but decreases to 0.7 to 0.9 µm.
There is an absorption band around 0.8–
1.4 µm and 2 µm. This is consistent withthe absorption bands of pyroxene andolivine.Often outside of the asteroid belt.

Ordinary chondrites

ReflectanceHowever, reflectance on the
short wavelength side islowered due to space
weathering.

LL chondrites
S-type asteroid (Itokawa)
Fig. 5.7: Comparison of spectrafrom S-type asteroids andordinary chondrites

WavelengthSource: The Solar System and its
Planets, Nippon Hyoronsha

118

Features of each type (2/2)
Type
X

Spectral form
Reflectivity gently increases at 0.4–0.9µm.

Distribution

Correlation with meteorites

Exist throughout the asteroid belt Iron meteorite
Metamorphic Tagish Lake
meteorite

Those with low reflectance of –0.04 at
0.55 µm are also called P-type.
M-type have reflectance of –0.1.

Enstatite chondrites
Achondrite

E-type have reflectance of –0.4.

AubriteD

Reflectance sharply increases at 0.45–
0.9 µm.

Near the Jupiter Trojans

Tagish Lake meteorite

V

Reflectance increases at 0.4–0.7 µm,
and abruptly drops at 0.7–0.9 µm.
Absorption bands are observed around
0.8–1.4 µm and 2 µm (absorption bandsof pyroxene).

Several percent of asteroids inthe asteroid belt

Similar to HED meteorites,
which are basaltic meteorites.
Asteroid Vesta origin?

119

Meteorite classifications
Frequency on EarthRocky meteorites

Chondrites·····86§← Undifferentiated meteorites

Achondrites···· 8§Classified fromratios of silicate
and metallic ironcomponents

Stony-iron meteorites······················· 1§

Differentiated meteorites

Iron meteorites (meteoric iron)·············· 5§Chondrites

Enstatite chondrites (EH, EL)
Ordinary chondrites (H, L, LL)
Carbonaceous chondrites (CI, CM, CO, CV, CR, CK, CH)120

Classification of Meteorites
Metachondrites
Carbonaceous

CI
Ivuna
CM
MurchisonCO
OrnansCompoundchondrites	

Ordinary

Chondrites

L
Low metal Fe
LL
Very low m-Fe

Kakangari

Ungrouped

H
High metal Fe

Primitive rocky meteorites
containing chondrules
A>63@71/:5@/7<A	Achondrites

Rumuruti

CK
Karoonda*2 Howardites, 	

Eucrites,
Diogenites 	

CH
High metal
CB
BencubbinEL
Low metal Fe

1There are otherPrimitive1

Acapulcoites

Enstatite

Aubrites

categories in addition
EH
High metal Fe
to the two below	

Differentiated meteorites in which
nearly same amounts of iron-nickel
alloys and silicate minerals coexist	

Stony-Iron Meteorites
Lodranites
HED *2
Vesta origin

Angrites

Pallasites
Ureilites

Lunar Met.

CR
Renazzo

Enstatite

Chondrule-freedifferentiated rocky
meteorites	

Itokawa particles	

CV
Vigarano

Italics indicate namesof representative
meteorites	

Mars Met.

Mesosiderites

Considered as the origin of
metal cores in planetesimals	

Iron Meteorites%&%Asteroid naming Asteroid discoverer receives naming rights.
 Here, “discoverer” means the person making first observations allowing for estimation oforbit.

 When asteroids are discovered, they are assigned a provisional name.
 After a number of observations allowing for sufficiently precise determination of orbit,
asteroids are assigned a permanent designation.

 Names can be proposed after a permanent designation is assigned.
 Proposed names are approved by the International Astronomical Union’s Committee on
Small Body Nomenclature.

Naming requirements:
A pronounceable (preferably single) word of 16 or fewer characters
Names from political or military events or persons allowed only 100 years after occurrence (forpersons, 100 years after death)
Pet names are not allowed
Restrictions on names of asteroids in special orbits
Names similar to those of existing celestial bodies are not allowed
Names for advertising or commercial purposes are not allowed122

Asteroid provisional names
1999 JU3
Year ofdiscoveryMonth of discovery
Char.
A
C
E
G
J
L
N
P
R
T
V
X

Date
tt Char.
Date
Jan. 1–15
B
Jan. 16–31
Feb. 1–15
D
Feb. 16–29
Mar. 1–15
F
Mar. 16–31
Apr. 1–15
H
Apr. 16–30
May 1–15
K
May 16–31
June 1–15
M
June 16–30
July 1–15
O
July 16–31
Aug. 1–15
Q
Aug. 16–31
Sep. 1–15
S
Sep.16–30
Oct. 1–15
U
Oct. 16–31
Nov. 1–15
W
Nov. 16–30
Dec. 1–15
Y
Dec. 16–31

Note: “I” is not used

Order ofdiscoveryMonth ofdiscovery

Order of discovery
A = 1stD = 4thG = 7thK = 10thN = 13thQ = 16th
T = 19thW = 22nd
Z = 25th

B = 2ndE = 5thH = 8thL = 11thO = 14th
R = 17th
U = 20thX = 23rd

C = 3rdF = 6thJ = 9thM = 12th
P = 15thS = 18th
V = 21st
Y = 24th

Note: 26th discovery is A1, 27th is B1, 51st is
A2, etc.
Note: Subscripts are not used when illegibility would
result, or typesetting does not allow

123

8. Science%&(Science from an asteroid sample return
Comparison with other asteroids	

Molecular cloud	Primitive solar
system nebula	Component materials and internal structure	Positioning with regard to Itokawa	

LL chondrites?	

Space
Rubble pile?	

weathering	

Meteoroid impacts	

Crater formation	

Itokawa
Rubble origins?	

Solarsystem	

Hayabusa

Regolith science	Interactions with space environment	New dating indices	

In addition to sciencefrom Itokawa…
Organics,
water

Ryugu


Cosmic ray irradiation	

Solar wind exposure	

Surface exploration
through impact
experiment■ Elucidating the birth and evolution of the solar system

Solar system’s past· What substances existed, and in what state?
· How did the planets arise and evolve?
· What were the raw materials (organic matter, water) for life?

■ Calibration of meteorites

Solar system’spresent

(© JAXA) 	

 How are meteorites and asteroid samples related?
Note: Many meteorites have been collected, but they are contaminated by the Earth’s
atmosphere, water, etc., making it difficult to determine their situation when they were inspace. Through comparisons with asteroids, meteorites can become valuable samples.%&)

Science: Elucidating the birth and evolution of the solar
systemCrosssections

TopicsPrimitive solar system disk (gas + dust)

Planetoids
Dust_Planetoid formation

` Investigating the materials thatformed the planets
What materials existed in theprimitive solar system disk, andhow did they change up to theformation of planets?

Primitive planets
Coalescence of planets

Terrestrial planet formation

a Investigating the formation
processes of the planets
How do celestial bodies growfrom planetoids to planets?

Gas giant formation

Loss of solar disk, completionof solar system(© JAXA) 	

%&*` Investigating the materials that formed theplanets The universe is thought to have been created 13.8 billion years ago. Following that, stellar evolution

produced various elements, which were scattered into space. The solar system was formed from thesematerials approximately 4.6 billion years ago, but the materials that were present in space at that timeremain unknown.

 We will clarify the distribution of substances in the original solar system disk.
 We will clarify how these substances changed on celestial bodies after their initial formation.
tttttt↓Finally, we will elucidate the materials that formed the planets, oceans, and life.

Keywords:
 Pre-solar particles: Particles from the interstellar molecular cloud brought into the solar system
 Calcium–aluminum-rich inclusions (CAI): Substances that record high-temperature states in theearly solar system Interactions between minerals, water, and organics: Diversification of organic matter on earlycelestial bodies
 Thermal metamorphism, space weathering: Material changes occurring within or on the surface ofcelestial bodies after their formation%&+Elucidation of organics by Hayabusa2
Volatile substances, such as water and organic matter, form on dust surfaces in molecular
clouds. It is thought that these change due to aqueous metamorphism and thermal denaturationin primitive solar system discs and planetoids, eventually accumulating on Earth and providing
materials for life. We will clarify what kinds of substance existed during this process.

Planetoid

Molecular cloud →tPrimitive solar disc

Accumulation

Metamorphosis

Transition of carbon materials on dust surfaces

EarthInteractions between materials, water, and organics
Minerals

Organics
Ice
Water in minerals

Asteroid

Chirality of amino acids

Earth
OceansLife
Left-handed (L-configuration) andright-handed (D-configuration)amino acids

Life on Earthalmostexclusively
uses lefthanded aminoacids. But
why?

%&,

Reference: Molecules containing carbon found in space(From Sakai (2011), Yuseijin)

Topics for Hayabusa 2What kind of organics were present in the primitive solar
disc before the creation of Earth?

%&-a Investigating planetary formation
Rubble pile body
Dust

(Minerals, water,organics)

Planetoids
ImpactMetal
core

PlanetDifferentiated body Elucidate the structure of planetoids that eventually became planets.
 Elucidate what processes occurred during the collisions, coalescence, and accumulationof celestial bodies.
tttt↓
tElucidate formation processes from planetoid to planet

Keywords:
 Rubble pile bodyª A celestial body formed from accumulated rubble
 Impacts and coalescence: When celestial bodies collide, the resulting fragments can
combine to form a new body Accumulation: Accumulation of fragments resulting from a collision via the force ofgravity%'$Reference: History of the sun, planets, and Earth
We will investigate events in the early solar system asrecorded in asteroids, thereby elucidating the process bywhich planets formed.

Gas
Water ice
Organics
Silicate dust
etc.

Appearance of main sequencestars: –10 Myr.

Adhesion / destruction
Evaporation/ condensation /melting
Metamorphism / alteration
Material mixing and circulationetc.

Accumulation into central
star / dissipation outside
the system
Collision / destruction / reaccumulation
Metamorphism / melting
differentiation by (parent body)internal heat generation

Galactic cosmic rays
Solar radiation,
windSpace weathering,
surface thermal
metamorphism

Near-EarthasteroidCollision, destruction,
orbit alteration

Life (–10 Myr.)

Meteorites,
dust

Missing Zone (planetoid=virtual body)
CAI formationCollision,
destruction,
accumulation

0.2Ma
First humans

Birth of life?

3800MaOldest sedimentary rockBirth of ocean & life?

Present

Magma oceansCM chondrite carbonate
minerals

Accumulation ofprimordial Earth

4563MaWater, organics
supply to Earth

Primordial planet
formation

4564MaHED differentiation

Allende chondrule

4568Ma 4565Ma
CAI solidification

Earth

Elucidation of
the Missing
Zone

Aqueous metamorphism (–200 ])Thermal metamorphism (–800 ])Chondrule formation (Ma«1 million years)

Main sequence
star

WTTS

Asteroid
accumulation

CTTS

Planetoid

Formation of primordial solar
discPrimordial star

Dispersion of discgas

Bipolar molecular flow

(virtual body)

Accumulation of gas
and dust from molecular
cloud coreContraction of molecular cloud
core, star formation

Planetoids

SunHayabusa&
Hayabusa2%'%9. International cooperation%'&Overview of international cooperation for
Hayabusa 2
USA

NASA
OSIRIS-REx(101955) Bennu

EuropeDLR
CNES
133

EU/US sample return missions1999 JU3(Ryugu)

EU

U.S.

Japan

Bennu

2014tlaunch2008 EV5

2016tlaunch(by JPL) 	

(by M. Busch) 	

Proposed but
not selected(Feb 2014)

Cooperation

Hayabusa2 

OSIRIS-REx

Solar system origins

Marco Polo-R%'(Comparison of mission target asteroids(101955) Bennu
1999 RQ36
9KTTQP

(341843)
2008 EV5(65803) Didymos
1996 GT;>6=6>!=4¶

9CSEQ<QNQ!=

30=?

>CORNGSGUVSP>CORNGSGUVSPPQUTGNGEUGF 6ORCEU

?ZRG

1

2

BM

>K[G

(-&O

($$O

+,$O>JCRG

(by D. Lauretta) 	

<GSKQF("&-+J'"+&)J

0YKT=0#3GE.,+#!*)=0#3GE.%$)#!**

0NDGFQ

$"$(*

$"%'+

:QUG

(by NASA) 	

(by M. Busch) 	

&"&)-'J$"%)DKPCSZ%');SDKUTQH=ZVIV 1GPPV &$$,4A)

Ryugu	

2008 EV5	

Bennu	%'*