III. HISTORY OF THE MISSION

Transcription

1 III. 0 Development phase 0 Dec Mar Dec Mar Apr May Nov Apr Launch Critical operations Initial function check Cruising phase Earth swing-by Southern hemisphere station operations Phase- ion engine operation Phase- ion engine operation Launch Rocket - H-IIA-6 (type 0) Planned launch date - 0 Nov 0 ::8 (Delayed due to weather) Actual launch date - Dec 0 ::0 Possible launch window - 0 Nov~ Dec 0 Launch location - Tanegashima Space Center Sub-payloads accompanying launch - Shin en (Kyushu Institute of Technology) - ARTSAT-DESPATCH (Tama Art University) - PROCYON (co-research by University of Tokyo and JAXA) Critical operations - Solar array panel deployment, sun acquisition control - Sampling device horn extension - Release launch lock on the retaining mechanism for the gimbal that controls ion engine direction - Confirm spacecraft tri-axial attitude control functions - Ground-based confirmation of functions for precise trajectory determination system Initial functional confirmation - Confirmation of ion engine, communications, power supply, attitude control, observation devices, etc. - Precise trajectory determination 08 Jan Jun 0 7 Phase- ion engine operation Asteroid arrival

2 H-IIA Launch VEHICLE 50m Satellite fairing m Satellite fairing (type S) Hayabusa HA0 [Standard] H-IIA naming: HA st/nd stage / number of LRB / number of SRB-A Length - 5 m nd Stage - st Stage - Mass - 8 ton SRB-A - SSB - 0m 0m 0m 0m Stage 7m Stage m Solid fuel rocket boosters 5m Co-payloads () Stage liquid hydrogen tank Stage liquid oxygen tank Stage engine Stage liquid oxygen tank Stage liquid hydrogen tank Solid rocket booster Rocket flight plan Altitude (km) Inertial velocity(km/s) 0:0:0 Liftoff :: Solid rocket booster burn completes # 6.6 0::8 separates # 5 0::0 Satellite fairing separation 7.8 0:6:6 Stage engine burn stop (MEC0) :6: Stage / separation 07 0:6:50 Stage primary engine start (SEIG) 0 0::8 stop (SEC0) :: Stage secondary engine start (SEIG) 50 :: stop (SEC0).8 :7:5 Hayabusa separation 88. :5:55 Shin en separation :58:5 ARTSAT-DESPATCH separation 8.7 ::5 PROCYON separation # At burn chamber max. pressure % # Thrust strut cutoff 0m Stage engine Satellite fairing separation Primary engine burn stops (MEC0) -0 Stage primary engine stop (SEC0) Stage secondary engine start (SEIG) Geodetic latitude [northern latitude, deg] Stage secondary engine stop (SEC0) Hayabusa separation Shin en separation ARTSAT-DESPATCH separation PROCYON separation Geodetic longitude [eastern longitude, deg]

3 Initial function check (details) Jan Functional confirmation of X-band mid-gain antenna beam pattern measurements, acquisition of actual data, and X-band communication equipment - Power system (battery) function check 0 - Near-infrared spectrometer (NIRS) inspection - Inspection of thermal infrared camera (TIR), deployable camera (DCAM), Optical Navigation Camera (ONC) -5 - Function check for attitude and trajectory system (all devices) 6 - Inspection of miniature rover (MINERVA-II) and lander (MASOT) 7 - Inspection of re-entry capsule and impactor (SCI) 8-5-point pointing test of X-band high-gain antenna (XHGA), pre-operation of ion engine - - Ion engine baking -6 - Ion engine test operation (ignition) *performed for each engine [/ : ion engine A; / ion engine B; /5 : ion engine C; /6 : ion engine D] 7 - Precise trajectory determination, Delta Differential One-way Ranging (DDOR) [No operations on /8, / ] Ka-band communications device actual data acquisition, antenna pattern measurements -0 - Ka-band DSN station DOR, lensing tests - Ion engine pre-operations -5 - Ion engine paired test operations [/: A+C; /: C+D; /: A+D; /5: A+C] 6 - Ion engine tri-set testing: A+C+D -0 - Paired engine -hour continuous autonomous operation: A+D - Function check of laser altimeter (LIDAR), laser range-finder (LRF), flash lamp (FLA) 0-/ - 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 May Jun Sep Oct Dec Initial operations phase complete, followed by normal operations phase. EDVEGA phase- IES operation Solar sail mode operations (maintains fuel-free solar orientation using only RW out of. Other RWs are kept in the OFF state) Three IES operate in -hour mode ( ITR-A+C+D ) EDVEGA phase- IES operation Solar sail mode operation starts IES-TCM ( precise trajectory control for swing-by ) Precise guidance phase ( TCM by RCS twice ) Earth swing-by ~06//E Southern hemisphere station operations ( by DSN Canberra and ESA Malargüe only ) 5

4 Phase- ion engine operations 0 Jan 08~ Jun 08 Phase- ion engine operations Nov 06~6 Apr 07 Arrive at 7 Jun 08 Launch Dec 0 Earth swing-by Dec 05 Phase- ion engine operations Mar 06~ May (incl. added burns) Trajectory to Hayabusa orbit Earth orbit Mar May Jun Jul Aug Oct Nov , Transfer phase- ion engine operations start Transfer phase- 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 -axis attitude control wheel STT Mars observations (OPNAV practice) ONC fixed-star observations DSN-UDSC uplink transfer testing Transfer phase- ion engine operations start 05 Before swing-by After swing-by Units Accel. m/s Time H Initial functioning IES operations testing confirmation /~ IES powered flight 0 5/~ IES max. thrust test 6/~6 IES powered flight 0 /~ IES powered flight. /~5/ Phase - ion engine operations /~/6 Phase - ion engine operations /0~6/ Phase - ion engine operations 7 78 at times 5 5 at times 75 Apr May Jun Sep Nov Dec ONC-T imaging near L5 Transfer phase- ion engine operations end ONC imaging of Jupiter and fixed stars RCS autonomous maneuvering tests Reset internal clock (TI) to zero DSN-SSOC real-time Doppler transmission testing DSN-UDSC uplink transfer testing IES test maneuvers Jan Feb Jun Transfer phase- ion engine operations start First observations Transfer phase- ion engine operations end Asteroid approach navigation start Asteroid arrival 6

5 05 06 Description of primary operations Solar sail mode ( 05 ~ ) A new technology that requires only a single reaction wheel; no fuel needed A new technology for Hayabusa 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 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 months of the.5-year flight. Utilizes force (pressure) from sunlight for batteries, etc. Only one reaction wheel (RW-Z) left ON RW-X RW-Z RW-Z RW-Y Scientific results from the swing-by ( Dec 05 ) ONC-T Color Earth image Intensity distribution of light reflected from vegetation TIR TIR thermal image ONC-T color image LIDAR Successful laser reception at 6.7 million km (0.05 AU) on Dec 05 NIRS Light absorption by water molecules in Earth's atmosphere Mars imaging ( May - Jun 06 ) May, Jun 06 We performed observations, taking advantage of an alignment of Hayabusa, Earth, and Mars. (Observations by ONC-T, NIRS, TIR) Near alignment ONC-T image of Mars :6 May 06(JST) Distance (AU) Mars Venu Mercury Earth Hayabusa Distance (AU) 7

6 Uplink transfer ( Jun - Nov 06 ) Uplink transfer technology testing Jun 06 : between DSN stations Nov 06 : between Usuda DSN Previous method Uplink transfer Communications temporarily cut Station A Communications not cut Station A Station A Communications not cut Station B Station B Station B Ka-band Communications DDOR ( Jun - Jul 06 ) Ka-band technology testing : Jun 8 Jul 06 Jun Jul : Ka-band communications testing at DSN Stn (Goldstone) Success from approx. 50 million km! Jul : Ka-band DDOR testing between NASA ESA stations (NASA DSN : Goldstone, ESA : Malargüe) World-first Ka-band DDOR between organizations! 5 8 Jul : Ka-band communications testing at ESA station X-band (8 GHz) : Normal operations Ka-band ( GHz) : Can transmit approx. times more data than X-band. Used to send asteroid observation data to Earth. waves from the quasar 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. 07 ~ Imaging at L5 ( 8 Apr 07 ) Three sets of four continuous images at 0 min intervals from the Optical Navigation Camera ( ONC-T ) telescope Date : 8 Apr 07 (JST) Exposure time : 78 sec (longest exposure) Results : No moving objects were seen in any sets Sun Orbital direction Earth Asteroid arrival Earth Launch Loc. At Sun Earth L5 Sun Earth system Lagrange points L, L5 8

7 Jupiter observation ( 6-7 May 07 ) Date : 6 May 07 7:0 (UCT), 7 May 07 0:0 (JST) View angle : deg Exposure time : 0. s Wavelength : v-band (550nm) Distance to Jupiter (6 May 07 7:0 UT) :.8565 au ( 6.70 x 0 8 km ) Magnitude as seen from spacecraft : -. Imaging objective : Various devices aboard Hayabusa 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 07 ) Time (TI) reset of the spacecraft clock Clock is reset through operations on 5 Sep 07. No need for further resets until return to Earth. Spacecraft-internal time counter : bits Time count: count = approx. ms ( ms = /000 s), bits allows counting to,,67,6 (approx. yr 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 First observation of ( 6 Feb 08 ) Successful imaging of by the onboard ONC-T camera on 6 Feb 08 Observation conditions were good on this day; was in the ONC-T FoV without making large attitude corrections. Distance from spacecraft to was approx.. million km is moving in the direction of the pink arrow. Three images are overlaid. View angle in the image is 0.8 deg

8 IV. Trajectories Trajectories overview After launch, the spacecraft enters a trajectory close to Earth orbit, and returns to Earth for a swing-by exactly year later. After the swing-by, it enters a trajectory close to orbit of asteroid, arriving there after about two orbits. It will remain at over a little more than one revolution around the sun. Dec 0 - Launch Dec 05 - Earth swing-by 7 Jul 08 - Asteroid arrival Nov - Dec 0 - Asteroid departure Nov - Dec 00 - Return to Earth Arrival at 7 Jun 08 Hayabusa trajectory orbit Earth orbit Earth swing-by Dec 05 Launch Dec 0 After that, it will leave, revolve around the sun for a little more than one orbit, then return to Earth Launch Earth swing-by Launch near, return and Earth swingby near 6. There is little distance between Earth and Hayabusa. Earth swing-by First orbit After Earth swing-by near 7, Hayabusa leaves Earth and gradually approaches (at ) First second orbit (asteroid arrival) While making one more orbit from to, Hayabusa makes one more orbit while approaching. Stay at asteroid Hayabusa arrives at near 0, and travels with the asteroid for over one solar orbit to 8. Asteroid Earth Hayabusa departs at around 8, then heads directly to Earth to return the capsule near. Trajectories 0

9 Trajectories in rotational coordinates Dec 0 - Earth departure Dec 05 - Earth swingby 7 Jun 08 - arrival Nov Dec 0 - departure Nov Dec 00 - Earth re-entry C = km /s Ion engine total impulse = km/sec Re-entry speed =.6 km/s Total flight time = 6 yr (.5 yr cruising time) Total powered flight time =.5 yr Total flight distance = 5. billion km Earth swing-by Hayabusa approached Earth for a swing-by on Dec 05. Earth approach time: : 08 (JST) Passed approximately,00 km over the Hawaiian islands Swing-by trajectory Diagrams depicting orbits around the sun. These figures show orbits of Earth and Hayabusa around the sun. The degree of curvature of the Hayabusa orbit at the swing-by point thus appears small. リュウグウ出発 departure Nov Dec 0/- 0 Operational リュウグウ近傍 trajectory in 運用軌道 vicinity Trajectory リュウグウから for Earth 地球への復路軌 return from 道リュウグウの orbit 軌道 Earth 地球再突入 re-entry Nov Dec 00/- 00 Earth 地球スイングバイ swingby 05/ EDVEGA Dec 05 ループ loop Earth リュウグウ Arrival at 到着 Jun Jul 08/ 地球からリュウグ Transitionary trajectory ウへの遷移軌道 from Earth to Hayabusa Before swing-by (Sep 05) Overview of orbits Swing-by point Hayabusa trajectory orbit Earth orbit Arrival at 7 Jun 08 Launch Dec 0 Hayabusa Earth swing-by Dec 05 After swing-by (Jun 06) Trajectories

10 Primary operations before and after Earth swing-by Trajectories at closest Earth approach TCM / Hayabusa trajectory Sunward direction 05 / / /0- / Earth and lunar imaging by thermal infrared camera TCM /6 / / TCM - Cancel / /6 / Earth and lunar imaging by optical navigation camera (telescope), / Closest Earth approach / thermal infrared camera, near-infrared spectrometer (swingby) / Lunar orbit / / / Earth imaging by optical navigation LIDAR optical link camera (wide-angle), experiment / imaging tracking function check / / / Cancel Earth observation Earth imaging by optical navigation attitude and transition / camera (telescope) to cruise attitude and thermal infrared camera / Northern polar direction 8:58 JST Enter shad :08:07 JST Closest approac 0min Sunshade :8 JST :0 Leave shad 0:00 0:0 :00 :00 8:0 6:0 7:00 7:0 8:00 Hayabusa trajectory Sunward direction Solar escape velocity V E V e Earth region escape velocity (Earth at center) Change in velocity relative to sun Velocity V A becomes V E Velocity of Earth rotation V E V A Principle of swing-bys A method of changing the trajectory of the spacecraft by converting from potential energy to kinetic energy using the planet's gravity. Beside changing the traveling direction, it can also accelerate or decelerate. Approach velocity with respect to sun V A V a Approach velocity to Earth region (Earth at center) The direction of ball travel has changed by 0 deg, and its speed increased from 0 m/s to m/s. Ball 0m/s m/s 0m/s m/s m/s 0m/s Car Throw a ball at 0 m/s at a right angle toward a car travelling at 0 m/s. From the perspective of the driver, the ball approaches the car diagonally at approx. m/s. The driver catches the ball, and throws it at m/s in the direction of travel. The ball is now travelling at m/s with respect to the ground In this metaphor, Hayabusa is the ball, and Earth and its gravity are the car and the driver Trajectories

11 V. Near-asteroid operations 0 m Automatic / autonomous technologies GSP, GCP-NAV Guidance Sequence Program (GSP) - From sensor information, autonomous behavior patterns performed by the spacecraft can be efficiently rewritten and delivered from the ground. - 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 this information to determine risk assessments and how to handle emergency situations. Before starting autonomous operations, ground commands are sent to rewrite tables in the spacecraft. - Efficient rewriting and instruction mechanisms are important for accommodating spacecraft restrictions on communications capacity and computer memory. Altitude Actual position Predicted 0 km position Actual trajectory Release TM Estimation error sphere Initial Position for GCP-NAV ΔV ΔV Set-point trajectory Final correction ΔV by GCP-NAV Altitude limit of GCP-NAV Synchronize with asteroid's surface TM Target point Approach phase GCP-NAV Final descent phase 6DOF Control Sensors in Use Ground Control Point Navigation (GCP-NAV) - Used for remote operations during approach from 0 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. ONC(onboard navigation camera)-w LIDAR( bram laser sensor) LRF(bram laser sensor) TM/FLA(flash lamp) Spacecraft trajectory calculation near the asteroid Considering the forces described here, calculate the trajectory of the spacecraft. Because planetary orbits around the sun are well known Known EscapeΔV Earth Gravity Gravity Gravity Planet True Path Gravity Sun Solar radiation pressure Propulsion Estimated If gravity on can be estimated, we can learn its mass. Density can be calculated when volume is known by shape estimation Leaving HP. Starting GCP-NAV (Ground / Onboard Hybird Navigation) Reference Path Alt. 0km Entering Autonomous Mode Alt. 00km Deploying Target Marker Alt. 0km Aligning Attitude to Local Surface Touch Down Alt. 0km Near-asteroid operations

12 Impactor operations sequence SCI separation Explosion & impact Horizontal avoidance Impact observation Vertical avoidanc DCAM separation Impactor: Debris and ejecta avoidance Impactor operates from above the asteroid (alt. several hundred meters) Debris avoidance : Debris rising due to explosion of the onboard impactor are avoided behind the asteroid. High-speed ejecta avoidance : Also avoid high-speed ejecta produced by projectile impact in. Low-speed ejecta avoidance : Low-speed ejecta falling back onto orbit around the asteroid are avoided by maintaining sufficient distance. Low-speed ejecta attaining very high altitudes are sufficiently slow as to not have a large effect, and chances of collision are low. Impactor separation Safe zone behind the asteroid Debris avoidance Explosion and impac Explosio High-speed ejecta avoidance Return to home position Low-speed ejecta avoidance Pinpoint touchdown Altitude 00 GCP-NAV Lateral direction ΔV to follow the asteroid's surface Target Markers (TM) TM separate at an altitude of several tens of meters, and flash lamps intermittently illuminate TM while cameras image them. By comparing differences in images when flash lamps are lit and when they are not, we can accurately extract TM without effects from surface patterns or sunlight. Facing toward identified TM, descend to the asteroid while using laser altimeter information to determine attitude and distance to the surface. 6-degree-of-freedom (position+attitude) gas jet injection control with high target tracking while minimizing fuel consumption is also a key technology. Use of multiple TM We will touch down near the artificial crater, and attempt to retrieve samples from exposed areas. We expect the artificial crater to have a diameter of around several meters. By approaching the destination point based on clues from multiple sequential TM, we can perform the touchdown with higher precision (a pinpoint touchdown) ~0 Using one TM Crater Range of error with no other TM for use as clues TM# Range of error for descent to TM # TM Onboard descent velocity control 0 [cm/s] TM release ONC-W FOV Switch LIDAR 0 [cm/s] Descent velocity 0 [cm/s] 0 [cm/s] Descent velocity 0 [cm/s] 0 [cm/s] TMT(Taeget Marker Tracking) mode DOF control Using multiple TM Range of error for descent to TM # with TM # visible TM# Near-asteroid operations Free Fall starts TD detection TM# Range of error for descent to TM # with TM # visibl