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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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