the 3rd nano-satellite symposium · tohoku university started development of sail deployment...
TRANSCRIPT
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Toshinori Kuwahara*, Yoshihiro Tomioka, Yuta Tanabe, Masato Fukuyama, Yuji Sakamoto, Kazuya Yoshida,
Tohoku University, Japan
The 3rd Nano-Satellite Symposium Micro/Nano Satellite & Debris Issues
December 12, 2011
Contents
1. Background
2. Suggested sail deployment mechanism
3. Development status
4. Orbital analysis
5. Results and outlook
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#1:SPRITE-SAT (RISING-1)
Launch: Jan. 2009 (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 (650-1000nm)
High resolution stereo images of cumulonimbus
Terrestrial luminous events in upper atmosphere
TAMU-2
To be launched around 2013
Small satellite development at Tohoku University
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Tohoku University has experience of 50 kg small satellite development
(Design, Development, Test, Launch, Operation)
RISESAT Mission - Design Conditions
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System Specification
Launch configuration
After panel deployment
Typical Design Condition of Auxiliary-Launch Microsatellites
Design conditions
Mass: < 50kg
Size: <50cm x 50cm x 50cm
Design life time: > 2years
Orbit: Sun-Synchronous Orbit
Typical orbit for Earth observation satellites
Large ground coverage
Altitude: 500 ~ 900 km
Inclination: ~98°
LTDN/LTAN: 9:00h ~15:00h
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12:00h 11:00h
9:00h 15:00h
15º 30º
LTDN
N
Sun
Ground Track Angle toward the Sun
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
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.
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Suggested Sail Deployment Mechanism
Deploy thin film with convex tape spring.
Can be mounted on the satellite’s body surface or inside the body.
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Sail Deployment Mech.
Sat. Sat. Sat.
On surface Inside body Half inside body
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
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Identified four Types of Sail Deployment Mechanisms
Type A B C D
Sail Areas [㎡] 0.5×0.5 1.5×1.5 2.5×2.5 4.5×4.5
Structure Size [mm] Φ50×30 φ100×40 φ150×50 φ200×80 or φ250×60
Switches [mm] φ25×5 φ50×7 φ63×7 φ100×7
Mass [g] 130 410 800 1780
Application
500~800km [kg] 0-1 1-10 10-50 50-120
900km [kg] - 0-1 1-10 10-50
A B C D
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Preliminary Functional Verification
Functionality of the design was verified for several models
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150mm
50mm
Size: B (Φ100mmx40mm)
Size: A (Φ50mmx30mm)
CubeSat “RAIKO”
(ISS 2012)
MicroSat “RISING-2”
2013~
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
Initial orbit altitude: 800km
Initial orbit altitude: 900km 12
SSO
Constant rotational rate: 0.173 º/s
Duration of de-orbit
Orbital Analysis (2/5) – Influence of rotational rate
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Weathercock stability analysis
Influence of rotational rate
Weathercock stability
Assume offset between mass and aerodynamic centers (250mm)
In case 50kg with C, feasible below 500km
Effect of rotational rate (relative to inertial frame)
No more difference if more than about 0.1º/s
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
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Changes in orbital elements Orbit Analysis (3/5) – Radiation Pressure
Orbital Analysis (4/5) – Active de-orbiting
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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
Sun
Max. Area projection Min. Area projection
1 rotation of right ascension of ascending node
Orbital Analysis (5/5) – Possibility of higher orbits De-orbiting from GTO Utilization of higher orbit for
micro-satellites Initial orbit: GTO Altitude=36000 -
300km 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?
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Summary and Outlook Summary Tohoku University started development of sail deployment mechanism for de-
orbiting 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 2012. 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
Thank you very much for your kind attention.
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