III. HISTORY OF THE MISSION

2011

Development phase

2014

Dec 3 Launch 5 Critical operations Launch 6 Initial function check Rocket - H-IIA-26 (type 202) Planned launch date - 30 Nov 2014 13:24:48 (Delayed due to weather) 2015 Actual launch date - 3 Dec 2014 13:22:04 Possible launch window - 30 Nov~9 Dec 2014 Mar Cruising phase Launch location - 2 Sub-payloads accompanying launch Dec Earth swing-by - Shin‘en 2 (Kyushu Institute of Technology) 3 - ARTSAT2-DESPATCH (Tama Art University) Southern hemisphere 4 - PROCYON (co-research by University of Tokyo and JAXA） station operations 2016 Critical operations - Solar array panel deployment, sun acquisition control - Sampling device horn extension Mar - Release launch lock on the retaining mechanism 22 Phase-1 Apr ion engine operation for the gimbal that controls ion engine direction - Confirm tri-axial functions May 21 - Ground-based confirmation of functions for Nov precise trajectory determination system 22 Phase-2 ion engine operation Initial functional confirmation 2017 - Confirmation of ion engine, communications, power supply, attitude control, observation devices, etc. - Precise trajectory determination Apr 26

2018

Jan 10 Phase-3 Jun ion engine operation 3 27 Asteroid arrival

HISTORY OF THE MISSION 23 H-IIA Launch VEHICLE

H2A202 [Standard] 50m Satellite fairing (type 4S) H-IIA naming: H2A 1st/2nd stage / number of LRB / number of SRB-A Length - 53 m Mass - 289 ton Satellite fairing 2nd Stage - 1 SRB-A - 2 12m 2 1st Stage - 1 SSB - ｜

40m Co-payloads (3) Rocket flight plan Altitude Inertial (km) velocity(km/s) 0:0:0 Liftoff 0 0.4 Stage 2 Stage 2 liquid hydrogen tank #1 11m 0:1:39 Solid rocket booster burn completes 46 1.6 Stage 2 liquid oxygen tank #2 0:1:48 separates 53 30m Stage 2 engine 0:4:10 Satellite fairing separation 137 2.8 0:6:36 Stage 1 engine burn stop (MEC0) 202 5.6 Stage 1 liquid oxygen tank 0:6:44 Stage 1/2 separation 207 Stage 1 0:6:50 Stage 2 primary engine start (SEIG1) 210 37m 0:11:18 stop (SEC01) 254 7.8 20m 1:39:23 Stage 2 secondary engine start (SEIG2) 250 Solid fuel 1:43:24 stop (SEC02) 313 11.8 rocket boosters Stage 1 liquid Hayabusa 2 separation 15m hydrogen tank 1:47:15 889 11.4 1:53:55 Shin’en 2 separation 2867 10.4 1:58:5 ARTSAT2-DESPATCH separation 4418 9.7 10m 2:2:15 PROCYON separation 6068 9.2

Solid rocket booster #1At burn chamber max. pressure 2% #2 Thrust strut cutoff

0m Stage 1 engine 1 Satellite fairing separation 2 Primary engine burn stops (MEC0) 3 -90 Stage 2 primary engine stop (SEC01) 4 Stage 2 secondary engine start (SEIG2)

Geodetic latitudeGeodetic latitude, [northern deg] 5 -60 Stage 2 secondary engine stop (SEC02) 6 separation 7 Shin’en 2 separation 1 30 4 2 3 5 8 ARTSAT2-DESPATCH separation 6 9 PROCYON separation 0 7 8 9

-30

-60

-90 0 30 60 90 120 150 180 210 240 270 300 330 360 Geodetic longitude [eastern longitude, deg] HISTORY OF THE MISSION 24 Initial function check (details)

2014 12 7·8 - Functional confirmation of X-band mid-gain antenna beam pattern measurements, acquisition of actual data, and X-band communication equipment 9 - Power system (battery) function check 10 - Near-infrared spectrometer (NIRS3) inspection 11 - Inspection of thermal infrared camera (TIR), deployable camera (DCAM3), Optical Navigation Camera (ONC) 12-15 - Function check for attitude and trajectory system (all devices) 16 - Inspection of miniature rover (MINERVA-II) and lander (MASOT) 17 - Inspection of re-entry capsule and impactor (SCI) 18 - 5-point pointing test of X-band high-gain antenna (XHGA), pre-operation of ion engine 19-22 - Ion engine baking 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] 27 - Precise trajectory determination, Delta Differential One-way Ranging (DDOR) 2015 [No operations on 12/28, 1/1–2] Jan 4 5-7 - Ka-band communications device actual data acquisition, antenna pattern measurements 9-10 - Ka-band DSN station DOR, lensing tests 11 - Ion engine pre-operations 12-15 - Ion engine paired test operations [1/12: A+C; 1/13: C+D; 1/14: A+D; 1/15: A+C] 16 - Ion engine tri-set testing: A+C+D 19-20 - Paired engine 24-hour continuous autonomous operation: A+D 23 - Function check of laser altimeter (LIDAR), laser range-finder (LRF), flash lamp (FLA) 20-3/2 - 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

History of flight

Mar 2 Initial operations phase complete, followed by normal operations phase. 3 EDVEGA phase-1 IES operation 21 27 mode operations May (maintains fuel-free solar orientation using only 1 RW out of 4. Other RWs are kept in the OFF state) 7 12 Three IES operate in 24-hour mode ( ITR-A+C+D ） 13 Jun 2 EDVEGA phase-2 IES operation 6 9 Solar sail mode operation starts Sep 1 IES-TCM ( precise trajectory control for swing-by ) 2 Oct 1 Precise guidance phase ( TCM by RCS twice ) Dec 3 Earth swing-by ~2016/4/E Southern hemisphere station operations 2016 ( by DSN Canberra and ESA Malargüe only )

HISTORY OF THE MISSION 25 Phase-3 ion engine operations 10 Jan 2018~3 Jun 2018 Phase-2 ion engine operations 22 Nov 2016~26 Apr 2017

Arrive at Ryugu Trajectory to Ryugu 27 Jun 2018 Hayabusa 2 orbit Launch Earth orbit 3 Dec 2014 Earth swing-by 3 Dec 2015 Phase-1 ion engine operations 22 Mar 2016~21 May (incl. added burns)

2016 2015 Accel. Time

Before swing-by Units m/s H Mar Initial functioning IES operations testing Transfer phase-1 ion engine operations start 22 confirmation May Transfer phase-1 ion engine operations end 21 3/3~21 IES powered flight 1 2 44 409 24 Mars observations (–Z Mars orientation) Jun 5/12~13 IES max. thrust test 3 4 24 1 6/2~6 IES powered flight 2 2 11 102 9 IES powered flight 3 14 Light pressure confirmation operations 9/1~2 2 1.3 12 20 2016 22 DSN-DSN uplink transfer testing 23 After swing-by DSN Ka-band communication testing 3/2~5/21 Phase -1 ion engine 3 127 798 29 operations 2 at times Jul 3 11/22~4/26 Phase -2 ion engine 3 435 2593 5 ESA Ka-band compatibility testing operations 2 at times 8 Aug 3 Transition to attitude control solar sail mode Oct 8 Transition to 3-axis attitude control wheel 2018 STT Mars observations (OPNAV practice) 11 ➡ 16 1/10~6/3 Phase -3 ion engine 2 3 393 2475 operations 19 ONC fixed-star observations 22 Nov 2,4 DSN-UDSC uplink transfer testing 22 Transfer phase-2 ion engine operations start

2017

Apr 18 ONC-T imaging near L5 26 Transfer phase-2 ion engine operations end May 16 ONC imaging of and fixed stars 28 30 RCS autonomous maneuvering tests Jun 20172018 1 Sep Jan 5 Reset internal clock (TI) to zero 10 Transfer phase-3 ion engine operations start Nov Feb 18 DSN-SSOC real-time Doppler 26 First Ryugu observations 28 transmission testing Jun Dec 3 Transfer phase-3 ion engine operations end 2 DSN-UDSC uplink transfer testing Asteroid approach navigation start 26 IES test maneuvers 27 Asteroid arrival 27

HISTORY OF THE MISSION 26 2015 Description of 2016 primary operations

Solar sail mode ( 2015 ～ )

A new technology that requires only a single reaction wheel; no fuel needed ・A new technology for Hayabusa2 that utilizes findings from Hayabusa and IKAROS ・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. ・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.

Utilizes force (pressure) from sunlight for batteries, etc.

Only one reaction wheel (RW-Z1) left ON

RW-X RW-Z1 RW-Z2 RW-Y

Scientific results from the swing-by ( 3 Dec 2015 )

ONC-T Intensity distribution TIR LIDAR Color Earth image of light reflected TIR thermal image ONC-T color image Successful laser reception at 6.7 million km (0.045 AU) from vegetation on 19 Dec 2015

NIRS3 Light absorption by water molecules in Earth's atmosphere

Mars imaging ( May - Jun 2016 )

Near ・24 May, 1 – 9 Jun 2016 alignment ・We performed observations, taking advantage of an alignment of Hayabusa2, Earth, and Mars. (Observations by ONC-T, NIRS3, TIR) Distance (AU) Venu

Mercury Earth Hayabusa 2 Mars Ryugu ONC-T image of Mars 21:46 24 May 2016(JST) Distance (AU)

HISTORY OF THE MISSION 27 Uplink transfer ( Jun - Nov 2016 )

Uplink transfer technology testing Previous method Uplink transfer 22–23 Jun 2016 : between DSN stations 2–4 Nov 2016 : between Usuda–DSN

Communications Communications Communications temporarily cut not cut not cut Station A Station B Station A Station A Station B Station B

Ka-band Communications DDOR ( Jun - Jul 2016 )

Ka-band technology testing : 29 Jun–8 Jul 2016 waves from the quasar ・29 Jun–3 Jul : Ka-band communications testing at DSN Stn (Goldstone) ⬆ Success from approx. 50 million km! ・1–2 Jul : Ka-band DDOR testing between NASA–ESA stations (NASA DSN : Goldstone, ESA : Malargüe) ⬆ World-first Ka-band DDOR between 3 organizations! ・5–8 Jul : Ka-band communications testing at ESA station X-band (8 GHz) : Normal operations Ka-band (32 GHz) : Can transmit approx. 4 times more data than X-band. Used to send asteroid observation data to Earth.

DDOR：Delta 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 more ground 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.

2017 ~

Imaging at L5 ( 18 Apr 2017 )

Three sets of four continuous images at 30 min intervals from the Optical Navigation Camera ( ONC-T ) telescope Date : 18 Apr 2017 (JST) Exposure time : 178 sec (longest exposure) Results : No moving objects were seen in any sets Orbital direction Earth Asteroid Launch arrival Sun Earth Loc. At

Sun–Earth L5

Sun–Earth system Lagrange points L4, L5 HISTORY OF THE MISSION 28 Jupiter observation ( 16 -17 May 2017 )

Date : 16 May 2017 17:30 (UCT), 17 May 2017 02:30 (JST) View angle : 0.79 × 0.79 deg Exposure time : 0.1312 s Wavelength : v-band (550nm) Distance to Jupiter (16 May 2017 17:30 UT) : 4.48565 au ( 6.71044 x 108 km ) Magnitude as seen from spacecraft : -2.44 Imaging objective : 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. Jupiter as imaged by ONC-T

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. ・Spacecraft-internal time counter : 32 bits ・Time count: 1 count = approx. 31 ms (1 ms = 1/1000 s), 32 bits allows counting to 4,294,967,296 (approx. 4 yr 3 mo) ・Counter reverts to zero after reaching max value (like a car odometer) ・This is performed to avoid a counter value of zero during stay at Ryugu

First observation of Ryugu ( 26 Feb 2018 )

・Successful imaging of Ryugu by the onboard ONC-T camera on 26 Feb 2018 ・Observation conditions were good on 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

Ryugu is moving in the direction of the pink arrow.

Three images are overlaid. View angle in the image is 0.8 deg

HISTORY OF THE MISSION 29 I V. Trajectories

Trajectories overview Hayabusa 2 trajectory Ryugu orbit After launch, the spacecraft enters a trajectory close to Earth Earth orbit orbit, and returns to Earth for a swing-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 Launch remain at Ryugu over a little more than one revolution around Arrival at Ryugu 3 Dec 2014 the sun. 27 Jun 2018 3 Dec 2014 - Launch Earth swing-by 3 Dec 2015 - Earth swing-by 3 Dec 2015 27 Jul 2018 - Asteroid arrival Nov - Dec 2019 - Asteroid departure Nov - Dec 2020 - Return to Earth After that, it will leave Ryugu, revolve around the sun for a little more than one orbit, then return to Earth. ①① ⑦⑦ ⑫ ② ⑤ ⑥ ⑦ ② ⑥⑥ ⑧⑧ ⑫ ⑫ ③ ④ ⑪ ⑤ ⑤ ③ ⑨ ⑧ ⑨ ④ ⑪ ⑩ ⑪ ① ④ ③ ⑩ ⑨ ② ⑩

Launch → Earth swing-by Earth swing-by → First orbit Launch near ① , return and Earth swing- After Earth swing-by near ⑦ , Hayabusa2 by near ⑥ . There is little distance between leaves Earth and gradually approaches Earth and Hayabusa2. Ryugu (at ⑫ ).

⑳ ⑬ ⑳ ㉘ ㉘ ⑬ ⑬ ㉗㉗ ㉝ ㉓ ㉖㉖ ㉝ ⑭ ⑱ ㉝ ⑰ ⑲ ㉗ ㉙ ⑭ ⑲ ㉒ ㉑㉑ ㉔ ⑲ ㉘㉘ ㉘㉘ ㉜ ㉜㉜ ⑭ ㉕㉕ ⑯ ⑮ ⑱ ㉑ ㉚ ⑮ ⑱ ⑮ ⑳ ㉖ ㉛ ㉕ ㉒㉒ ㉔㉔ ㉙㉙ ㉛㉛ ⑯ ⑰ ⑯ ⑰ ㉓㉓ ㉚㉚

First → second orbit (asteroid arrival) Stay at asteroid Asteroid → Earth While making one more orbit from ⑲ to ① , Hayabusa2 arrives at Ryugu near ⑳ , and Hayabusa2 departs Ryugu at around ㉘ , Hayabusa2 makes one more orbit while travels with the asteroid for over one solar then heads directly to Earth to return the approaching Ryugu. orbit to ㉘ . capsule near ㉝ .

Trajectories 30 Trajectories in rotational coordinates Earth swing-by 3 Dec 2014 - Earth departure Hayabusa2 approached Earth for a swing-by on 3 Dec 2015. 3 Dec 2015 - Earth swingby Earth approach time: 19 : 08 (JST) 27 Jun 2018 - Ryugu arrival Passed approximately 3,090 km over the Hawaiian islands - Ryugu departure Nov–Dec 2019 Swing-by trajectory Nov–Dec 2020 - Earth re-entry Diagrams depicting orbits around the sun. These figures show 2 2 C3 = 21 km /s orbits of Earth and Hayabusa2 around the sun. The degree of Ion engine total impulse = 2 km/sec curvature of the Hayabusa 2 orbit at the swing-by point thus Re-entry speed = 11.6 km/s appears small. Total flight time = 6 yr (4.5 yr cruising time) Total powered flight time = 1.5 yr Total flight distance = 5.24 billion km

Ryuguリュウグウ出発 departure Nov–Dec2019/11- 122019 Trajectoryリュウグウから for Earth地球への復路軌 return from Ryugu道

リュウグウのRyugu orbit 軌道 Earth地球再突入 re-entry Operational Nov–Dec2020/11 -202012 リュウグウ近傍trajectory in Ryugu運用軌道 vicinity Earth地球スイングバイ swingby EDVEGA Dec2015/12 2015 EDVEGA ループloop Earth

Arrival at Ryugu リュウグウ 到着 地球Transitionaryからリュウグ Hayabusa2 Jun–Jul2018/6 2018-7 trajectoryウへの遷移軌道 from Earth to Ryugu

Before swing-by (Sep 2015)

Overview of orbits Swing-by point

Hayabusa 2 trajectory Ryugu orbit Earth orbit

Launch Arrival at Ryugu 3 Dec 2014 Hayabusa2 27 Jun 2018 Ryugu

Earth swing-by 3 Dec 2015 After swing-by (Jun 2016) Trajectories 31 Primary operations before and after Earth swing-by Trajectories at closest Earth approach

2015 TCM1 11/3 11/4 16:30 11/9 17:00 11/10-13 Northern polar 11/14 Earth and lunar imaging by direction 17:30 thermal infrared camera TCM2 11/26 11/19 Hayabusa18:00 2 trajectory Hayabusa 2 trajectory 11/24 18:58 JST 18:30 TCM3 - Cancel 11/26 Enter shad 12/1 11/29 Sunward Earth and lunar imaging by optical direction navigation camera (telescope), 12/3 thermal infrared camera, 19:08:07 JST 12/4 Closest approac 19:00 Closest Earth approach near-infrared spectrometer (swingby) 12/9 20min Sunshade Lunar orbit 12/3 19:30 12/14 19:18 JST Sunward 12/19 Earth imaging by optical navigation Leave shad LIDAR optical link camera (wide-angle), direction experiment 12/19 imaging tracking function check 20:00 12/22 12/4 Cancel Earth observation 12/24 Earth imaging by optical navigation 20:30 attitude and transition camera (telescope) to cruise attitude 12/29 and thermal infrared camera 21:00 1/3

Ve Earth region escape velocity (Earth at center) Solar escape velocity VE

Change in velocity relative to sun Velocity becomes VA VE

Velocity of Earth rotation

VE VA

Principle of swing-bys A method of changing the trajectory of the spacecraft by

converting from potential energy to kinetic energy using the Va Approach velocity to Earth region (Earth at center) planet's gravity. Beside changing the traveling direction, it can Approach velocity with respect to sun also accelerate or decelerate. VA

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

Ball 14m/s 10m/s 10m/s 14m/s 14m/s 10m/s Car Throw a ball at 10 m/s at a right From the perspective of the The driver catches the ball, The ball is now travelling at 24 angle toward a car travelling at driver, the ball approaches and throws it at 14 m/s in the m/s with respect to the ground 10 m/s. the car diagonally at approx. direction of travel. 14 m/s. In this metaphor, Hayabusa2 is the ball, and Earth and its gravity are the car and the driver

Trajectories 32 V. Near-asteroid operations

Automatic / autonomous Spacecraft trajectory calculation near the asteroid Considering the forces described here, calculate the trajectory of technologies GSP, GCP-NAV the spacecraft.

Earth Sun Gravity Guidance Sequence Program (GSP) Solar radiation Known Gravity - From sensor information, autonomous behavior patterns pressure performed by the spacecraft can be efficiently rewritten and Because planetary delivered from the ground. orbits around Propulsion Estimated - We first obtain asteroid information that can only be derived the sun are Gravity well known Gravity through proximal observations, such as its surface conditions and reflectivity. Operators on the ground analyze this Planet If gravity on Ryugu can be information to determine risk assessments and how to handle estimated, we can learn its Ryugu mass. emergency situations. Before starting autonomous operations, ➡ Density can be calculated ground commands are sent to rewrite tables in the spacecraft. when volume is known by shape estimation - Efficient rewriting and instruction mechanisms are important for accommodating spacecraft restrictions on communications capacity and computer memory. Alt. 20km Leaving HP. Starting GCP-NAV Altitude (Ground / Onboard Hybird Navigation) Reference Actual True Path position Estimation error sphere Sensors in Use Path Predicted Initial Position 20 km position for GCP-NAV GCP-NAV Approach phase

Actual Set-point trajectory trajectory Entering Autonomous Mode ΔV Alt. 100km

ΔV Final correction ΔV by GCP-NAV Deploying Target Marker Altitude limit of 40 m Release TM GCP-NAV laser sensor) bram LIDAR(1 6DOF Control Final descent phase EscapeΔV Alt. 30km Synchronize with Aligning Attitude to Local Surface ONC(onboard navigation camera)-W1 navigation ONC(onboard TM/FLA(flash lamp) asteroid's surface laser sensor) LRF(4bram Touch Down TM Alt. 0km Target point

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

Near-asteroid operations 33 Impactor operations sequence

Impactor: Debris and ejecta avoidance Impactor operates from above the asteroid (alt. several hundred meters) 1 Debris avoidance : Debris rising due to explosion of the SCI separation Horizontal Vertical avoidance avoidanc onboard impactor are avoided behind the asteroid. 2 High-speed ejecta avoidance : Also avoid high-speed ejecta produced by projectile impact in 1 . Impact 3 Low-speed ejecta avoidance : Low-speed ejecta falling back observation onto orbit around the asteroid are avoided by maintaining DCAM3 sufficient distance. Low-speed ejecta attaining very high separation altitudes are sufficiently slow as to not have a large effect, Explosion and chances of collision are low. & impact Impactor separation

Safe zone behind 1 Debris avoidance the asteroid

Ryugu 2 High-speed Explosio ejecta avoidance Explosion and impac

3 Low-speed ejecta avoidance Return to home Ryugu position

Altitude GCP-NAV

Pinpoint touchdown Lateral direction ΔV to follow 100 the asteroid's surface

Onboard descent velocity control Target Markers (TM) 50 10 [cm/s] ・TM separate at an altitude of several tens of meters, and flash Switch LIDAR → 20 [cm/s] lamps intermittently illuminate TM while cameras image 40 them. Descent velocity ・By comparing differences in images when flash lamps are lit 10 [cm/s] → 20 [cm/s] 30 TM release and when they are not, we can accurately extract TM without Descent velocity effects from surface patterns or sunlight. ONC-W1 20 [cm/s] → 10 [cm/s] FOV ・Facing toward identified TM, descend to the asteroid while TMT(Taeget Marker Tracking) using laser altimeter information to determine attitude and 15 mode DOF control distance to the surface. TM ・6-degree-of-freedom (position+attitude) gas jet injection control with high target tracking while minimizing fuel 5~10 Free Fall starts 1 TD detection consumption is also a key technology.

Use of multiple TM Using one TM Using multiple TM

・We will touch down near the artificial crater, and attempt to Range of error for descent to TM #2 retrieve samples from exposed areas. Crater with TM #1 visible TM#3 TM#1 ・We expect the artificial crater to have a diameter of around Range of error for Range of error TM#2 descent to TM #3 several meters. By approaching the destination point based with no other TM Range of error with TM #2 visibl for use as clues for descent to on clues from multiple sequential TM, we can perform the TM #1 touchdown with higher precision (a pinpoint touchdown). Near-asteroid operations 34