Toshinori Kuwahara*, Yoshihiro Tomioka, Yuta Tanabe, Masato Fukuyama, Yuji Sakamoto, Kazuya Yoshida, Tohoku University, Japan

Transcription

1 Toshinori Kuwahara*, Yoshihiro Tomioka, Yuta Tanabe, Masato Fukuyama, Yuji Sakamoto, Kazuya Yoshida, Tohoku University, Japan The 3 rd Nano-Satellite Symposium Micro/Nano Satellite & Debris Issues December 12,

2 Contents 1. Background 2. Suggested sail deployment mechanism 3. Development status 4. Orbital analysis 5. Results and outlook 2

3 Small satellite development at Tohoku University #1:SPRITE-SAT (RISING-1) Launch: Jan (H-IIA) Demonstration of Image acquisitions by mission camera Coarse attitude control Deployment of the boom TAMU: Tohoku-Ångström MEMS Unit #2:RISING-2 Completed (- Sep 2011) FM system integration and verification Software update Mission Multi-spectrum observation with a Liquid Crystal Tunable Filter ( nm) High resolution stereo images of cumulonimbus Terrestrial luminous events in upper atmosphere TAMU-2 To be launched around 2013 Tohoku University has experience of 50 kg small satellite development (Design, Development, Test, Launch, Operation) 3

4 RISESAT Mission - Design Conditions System Specification Launch configuration After panel deployment 4

5 Typical Design Condition of Auxiliary-Launch Microsatellites Design conditions Mass: < 50kg Sun Size: <50cm x 50cm x 50cm Design life time: > 2years Orbit: Sun-Synchronous Orbit Typical orbit for Earth observation satellites 15:00h LTDN 12:00h 15º 11:00h 30º 9:00h Large ground coverage Altitude: 500 ~ 900 km Inclination: ~98 LTDN/LTAN: 9:00h ~15:00h N Angle toward the Sun Ground Track 5

6 Suggested Sail Deployment Mechanism World debris prevention activities: Inter-Agency Space Debris Coordination Committee (2002~) United Nations General Assembly: Committee on the Peaceful Uses of Outer Space (2007~) Limit the long-term presence of spacecraft and launch vehicle orbital stages in the low-earth orbit region after the end of their mission Purpose: Prevent satellites from becoming space debris after their operation in order not to disturb future new satellites/spacecrafts. Concept: Deploy a large sail triggered by electrical switch via commands just before the satellite terminates its mission life so that the area-to-mass ratio becomes large enough for de-orbiting. Utilizes atmospheric drag in order to decrease the orbit altitude/orbital energy to let the satellite re-enter into dense Earth atmosphere. Realize de-orbiting within 25 years after the activation of the mechanism Assumption: Small satellites burn out during the re-entry phase into dense Earth atmosphere and there is no risk for human activity on the Earth. Applied to orbits where Earth atmosphere practically still exists. 6

7 Requirements of Sail Deployment Mechanism for De-orbiting Fundamental requirements: Large enough area size for de-orbiting target space debris/satellites Light-weight in order not to disturb original mission objectives of the spacecraft Small size with effective/dense storing method High reliability Can survive in space environment for about 25 years of operation Can keep the form of the sail after deployment Easy to install into spacecraft structure Additional requirements: Prevent utilization of pyrotechnic If possible 3 dimensional sail is desired Low cost Passive deployment / no electrical motor or such. 7

8 Suggested Sail Deployment Mechanism Deploy thin film with convex tape spring. Can be mounted on the satellite s body surface or inside the body. Sail Deployment Mech. Sat. Sat. Sat. On surface Half inside body Inside body 8

9 Definition of Mass Category and Related Model Size Definition of mass category of small satellites Definition of required sail area for each mass category Target sizes of sail deployment mechanism 9

10 Identified four Types of Sail Deployment Mechanisms A B C D Type A B C D Sail Areas [ m2 ] Structure Size [mm] Φ50 30 φ φ φ or φ Switches [mm] φ25 5 φ50 7 φ63 7 φ100 7 Application Mass [g] ~800km [kg] km [kg]

11 Preliminary Functional Verification Functionality of the design was verified for several models CubeSat RAIKO (ISS 2012) Size: A (Φ50mmx30mm) 50mm MicroSat RISING ~ Size: B (Φ100mmx40mm) 150mm 11

12 Orbital Analysis (1/5) Influence of area-to-mass ratio Parameters Original orbit (at the sail deployment) Attitude, rotational rate Area-to-mass ratio Form of the sail (2D,3D) Atmospheric drag Gravity field Solar radiation pressure Duration of de-orbit depends on the initial conditions and mathematical models of above effects. Area of sail is set in safe-side Duration of de-orbit Initial orbit altitude: 800km SSO Constant rotational rate: º/s Initial orbit altitude: 900km 12

13 Orbital Analysis (2/5) Influence of rotational rate Effect of rotational rate (relative to inertial frame) No more difference if more than about 0.1º/s Influence of rotational rate Weathercock stability Assume offset between mass and aerodynamic centers (250mm) In case 50kg with C, feasible below 500km Weathercock stability analysis 13

14 Orbit Analysis (3/5) Radiation Pressure Changes in orbital elements Effect of solar radiation pressure Initial orbit: SSO with LTDN=12:00, Altitude=500km 3 different size of sail: 4.5m x 4.5m, 7.5m x 7.5m, 10m x 10m Observed changes in orbital elements The sail needs to be considerably large enough to produce meaningful effects Orbit altitude under the effect of solar radiation pressure 14

15 Orbital Analysis (4/5) Active de-orbiting Active de-orbiting with attitude control of the sail Initial orbit: SSO with LTDN=12:00, Altitude=500km Effect of atmospheric drag neglected Sail size: 10m x 10m Switch sail projection area toward the Sun between Max. and Min. Effective de-orbiting about 70km in each rotational period of the right ascension of ascending node Effective also in high-altitude orbits Max. Area projection Sun 1 rotation of right ascension of ascending node Min. Area projection 15

16 Orbital Analysis (5/5) Possibility of higher orbits De-orbiting from GTO Utilization of higher orbit for micro-satellites Initial orbit: GTO Altitude= km Sail size: 10m x 10m Radiation effects neglected Altitude of apogee: decreases Altitude of perigee: remains Low Earth Orbit region The altitude of apogee can be decreased down to the LEO region in about 2000 days. Suggested de-orbit mechanism also works for micro-satellites in GTO High-altitude orbit can be utilized for micro-satellites? 16

17 Summary and Outlook Summary Tohoku University started development of sail deployment mechanism for deorbiting of small satellites. A functional model was developed and its functionality was evaluated. Tohoku University is now developing different sizes of sail deployment mechanism which are going to be demonstrated on microsatellites in near future. This mechanism enables active prevention and reduction of space debris. Outlook Conduct environmental tests Mechanical, thermal vacuum, AO Continue orbital analysis for establishing effective utilization method of sail deployment mechanism. Size C will be developed by March Possibly can be standardized for future small satellite Investigate feasibility of applying to larger satellites ( >150kg ) Investigate feasibility of launching small satellites into higher altitude orbits such as MEO or GTO. 17

18 Thank you very much for your kind attention. 18