Abstract—The attitude dynamic module of LAPAN-ITB satellite simulator has been used in choosing the optimal Earth observation micro-satellite design, based on the attitude control subsystem power

consumption. In this research, further design choice to improve satellite performance is considered by deploying the satellite’s solar panels. Such configuration improves the power production capacity, but requires the satellite attitude control to point its solar panels side to the Sun to ensure the electrical supply to the batteries. The power consumption from the additional pointing requirements is calculated using the simulator, where 2 attitude control modes are selected, i.e. zero angular momentum and angular momentum bias. The results shows that the deployable panels increase the coverage performance of

the satellite by twice, compared to body mounted configuration, for both attitude control modes. The research has shown the effectiveness of LAPAN-ITB satellite simulator as design optimization tools.

Index Terms— Attitude Dynamic Simulation, Design

Optimization, High Fidelity Model, Micro-Satellite Design.

I. INTRODUCTION

The recent trend in satellite design optimization is using

optima search algorithm and high fidelity design model, with

optimization objectives in system level performance [1][2].

Such design optimization mode is also done by the astronautics

group of ITB, by performing an integrated preliminary and

detail satellite system design optimization, that involve search

algorithm and high fidelity model [3][4]. In the study, 3 system

performance parameters of Earth observation satellite, i.e.

image’s resolution, swath, and length/coverage were used as optimization objectives. The first two performance parameters

were optimized in the preliminary design phase. The result is a

pareto of satellite design choices that comply with the

determined design constraints. The last performance parameter

was optimized in detail design phase using satellite dynamic

simulator. The method has successfully selected the best

satellite configuration, even though most of the improvement in

the satellite image length is due to more precise attitude control

power consumption calculation, and not from the configuration

differences (i.e. satellite moments of inertia).

The objective of this research is to find more optimum

design, compared to the one perfomed in [4], using the satellite

Manuscript received May 9, 2016. This work was supported in part by ITB

Reseacrh Grant no. FTMD.PN-6-03-2016. Robertus H. Triharjanto is graduate student at Faculty of Mechanical and

Aerospace Engineering, Institut Teknologi Bandung (ITB), Indonesia.

Ridanto E. Poetro is lecturer at Faculty of Mechanical and Aerospace

Engineering, ITB, Indonesia.

Hari Muhammad is lecturer and head of Flight Physics Research Group at

Faculty of Mechanical and Aerospace Engineering, ITB, Indonesia.

Soewarto Hardhienata is lecturer at faculty of Mathematics and Natural

Science, Universitas Pakuan, Bogor, Indonesia

simulator. Here means that the design could increase the system

performance of the satellite significantly, without violating the

basic design constrains. In the previous satellite design

optimization [4], the image length/coverage is limited by the

power available for payload operation. Design option to

increase the power capacity in Earth observation

micro-satellite, among other, is by using deployable solar

panel. German’s micro-satellite, BIRD, Korean’s

micro-satellite, KITSAT-3, and Japan’s micro-satellite, Hodoyoshi-4, are some of the Earth observation

micro-satellites that use such strategy. BIRD satellite needs to

accommodate high power consumption since its thermal

infrared camera requires active cooling system [5]. KITSAT-3,

in addition of carrying 3-band medium resolution imager and

multiple scientific payloads, was designed to increase the

satellite bus power to 5 times its predecessor [6]. Meanwhile,

Hodoyoshi-4 is micro satellite with the highest ever downlink

data rate, which is used to transmit among other its 4-band high

resolution images [7]. All of those reasons need higher power

production capacity than what micro-satellite with body-mounted solar panels can provide. The 3 aforementioned

satellites have 2 solar panels deployed, in this case, in –Z

direction or away from nadir/camera direction.

In satellite with deployable solar panels, its attitude control

has to work almost all the time to ensure sun pointing to its

solar panel side. Failure in attitude control system could cause

the batteries do not have enough power supply to operate the

satellite. Such incidence was experienced by BIRD [8]. In

KITSAT-3, the sun pointing control mode becomes the default

(safe) mode. Therefore, after all other mode is switched-off

(intentionally or due to any failure) the sun pointing is selected.

From the attitude sensor components on-board, mostly the sun pointing mode is done using sun sensor and horizon sensor

[6][9].

LAPAN-ITB satellite simulator is developed by the lecturer

and students of FTMD, ITB, with collaboration with Center for

satellite Technology, LAPAN. The current version of attitude

dynamics module simulates satellite with 3 orthogonal reaction

wheels [10]. The simulator may calculate the effect of

deployable solar panel on the control torque that the reaction

wheels have to provide, which then can be used to calculate the

power consumption of the satellite attitude control subsystem.

II. THE DESIGN MODEL

A. Satellite Configuration

The design model of Earth observation micro-satellite used

in the research derived from previous the work on satellite

design optimization, which involve satellite attitude dynamic

simulator as a high fidelity model [4]. The study yield that the

Evaluating Deployable Solar Panel Option in Earth

Observation Micro-satellite Design Optimization

Robertus H. Triharjanto, Ridanto E. Poetro, Hari Muhammad, and Soewarto Hardhienata

Int'l Journal of Advances in Mechanical & Automobile Engg. (IJAMAE) Vol. 3, Issue 1(2016) ISSN 2349-1485 EISSN 2349-1493

http://dx.doi.org/10.15242/IJAMAE.IAE0716207 105

optimum satellite design has dimension of 60 x 60 x 55,0 cm,

which may produce 4-band multispectral image with resolution

of 13 m and swath of 200 km, for 34 minutes per orbit (or cover

the area of 2,85 million km2). The 2nd best configuration is

satellite with dimension of 60 x 60 x 55,9 cm, which may

produce 4-band multispectral image with resolution of 7 m and swath of 75 km, also for 34 minutes per orbit (or cover the area

of 1,07 million km2).

The limitation in the payload operation (capturing images

and send them to groundstation) in the above satellite

configurations is due to the limited power supply from the body

mounted solar panel. In such configuration, the effective solar

panel area exposed to the sun is equivalent to area of 1 side only

(see Fig. 1). Even with the best solar cell (GaAs triple junction),

the satellite need to hibernate after 34 minutes of payload

operation per-orbit to charge its batteries. Deploying the panel

on X+ and X- sides will increase the solar panel area exposed to

sun by three folds (see Fig. 2).

Fig. 1. Micro-satellite with body mounted solar panel (payload camera shown)

Fig. 2. Micro-satellite with deployable solar panel

Different than the design of BIRD, KITSAT-3, and

Hodoyoshi-4, the deployed solar panel the satellite model in

this research is on the Y+ side, or perpendicular with the

camera axis. By assuming that the satellite is launched as an

auxilary payload to Sun synchronous orbit, with flight path

from North to South and about 8-9 AM equatorial crossing,

the Sun vector will be at 45-60o incline East from the orbital

plane (see Fig. 3). Therefore, the deployed solar panel in Y+

side could acquire about 70% of the solar flux, when the

satellite is nadir pointing.

Table 1 contains, among others, the satellite model mass

properties. The properties of configuration 2 and 5 is taken

from the previous works [3][4]. The weight of the satellites of

configuration 2D and 5D (D = deployed) are calculated with

assumption that the structure to support deployable solar panel

is made from light materials (isogrid plate or carbon

composite), and the deployable mechanism is composed of

spring-hinges and latches. Using the geometry from Fig. 2, and

weight of the deployed panels, the inertia of the satellite models can be calculated. Considering that the change in weight is less

than 2%, and the change in moment of inertia is about 7% in X

& Z axis, and 26% in Y axis, the effect of the deployable panels

in the cross product of inertia in this calculation is neglected

(same as body mounted ones).

The power productions for such configurations are

calculated at 70% of the maximum capacity, since assuming

that the satellite maintain nadir pointing, the sun pointing is at

45-60o angle.

TABLE 1 MOMENT INERTIA, WEIGHT AND POWER PRODUCTION CAPACITIY OF THE

SATELLITE MODELS

Config. Ixx (kg.m

2)

Iyy (kg.m

2)

Izz (kg.m

2)

Weight (kg)

Power prod. /orbit (Whr)

2 4,28 4,28 4,58 76,4 137,6

5 4,20 4,20 4,56 76,0 135,3

2D 4,62 5,41 4,90 79,0 289,0

5D 4,53 5,30 4,87 78,6 284,1

B. Attitude Control Strategy

For satellite with deployable solar panels, the attitude

requirement is to have the panels points to the Sun when not in

eclipse. In this design case, it means +Y pointed to the Sun. For

image acquisition, additional requirement is needed, i.e. the Z+

to be nadir pointed. When the satellite is in eclipse, no attitude

control is required. KITSAT-3 has 4 attitude control mode. When charging its

battery, it used pure sun pointing for its solar panel side. When

taking images, it used nadir pointing mode for the camera. In

eclipse, the satellite has 2 attitude control modes, i.e. +Z toward

space when activating its science payload, and nadir pointing

when transmitting data to groundstation [9]. Such strategy

might put heavy load on the attitude control subsystem, since

there will be transient phase every time the control mode is

changed, as has been simulated in the previous research [4].

Therefore, the attitude control strategy chosen in this research

uses minimum number of modes, to avoid high consumption

during mode changing. In this research, 2 attitude control strategies are selected for

the satellite operation, i.e. zero angular momentum mode and

Y

X

Z

Y

Z

Int'l Journal of Advances in Mechanical & Automobile Engg. (IJAMAE) Vol. 3, Issue 1(2016) ISSN 2349-1485 EISSN 2349-1493

http://dx.doi.org/10.15242/IJAMAE.IAE0716207 106

angular momentum vector management mode. In zero angular

momentum mode, the satellite angular momentum is not

exactly zero. Instead, the satellite is set to have minimum

angular momentum so that its attitude control actuators

(reaction wheels) may work in good stability. The satellite

design use data from LAPAN satellite’s reaction wheels (IRE 303), which the rotation speed of no less than 200 rpm. With

this minimum angular momentum, the 3 reaction wheels will

always operate to keep the satellite +Z nadir pointing and +Y

face the East.

In this scenario the angular momentum will slowly increase

due to external torque disturbances, such as Earth magnetic

field, and gravity gradient. These may increase the angular

momentum, and therefore the compensation torque that the

reaction wheels have to provide, and power consumption will

increase. When the satellite’s angular momentum reached

certain level of reaction wheels absorption capacity, it will be

dumped using magneto-torquer. Since the satellite cross product of inertia only about 2% of

its major inertia, the angular momentum bias mode may be used

as the 2nd operation scenario (only has small nutation). In this

mode, the satellite maintain angular momentum vector at its Y

axis, which is the major inertia axis. Such strategy is once tried

in LAPAN-TUBSAT operation, with the magnitude of the

angular momentum is equivalent to 86% of the reaction wheel

absorption capacity [11]. To reduce power consumption, during

hibernation mode, only 70% of the angular momentum is

absorbed by Y axis reaction wheel, and therefore the satellite

rotated at rate of 1 deg/s in opposite direction of the wheel. In order for the solar panel to have good exposure to the Sun,

the angular momentum vector is maintained perpendicular of

the flight path/orbital plane, with +Y pointed Eastward.

Attitude control experiment in LAPAN-TUBSAT shows that in

such magnitude and direction, the angular momentum vector

only drifted by 1 deg/day, and can be compensated by daily

execution of the magneto-torquers [12].

Fig. 3. Attitude control scenario for angular momentum bias mode

Fig. 3 shows the imaging operation in the angular

momentum bias mode. It is performed by terminating the

satellite pitch (reaction wheel to absorb all angular momentum), then rotate in Y axis until nadir pointing is found

(by horizon sensor). The advantage of using this mode is that

only one reaction wheel operates at hibernation mode.

C. Modeling Implementation

The main governing equations in the attitude dynamics

module of LAPAN-ITB satellite simulator are eq. (1) and (2).

Based on the attitude control strategy mentioned in previous

subchapter, the difference with the problem implemented in [4]

is that system angular momentum (Hi) is not equal to zero. The

angular momentum is generated in (or dissipated from) the

system using magneto-torquers as the source of external torque,

modeled as the last term in eq. (2).

*

+

*

( )

( )

( )

+ [

] *

( )

( )

( )

+ [

] (1)

*

+

*

( )

( )

( )

+ *

( )

( )

( )

+

(*

+ *

( )

( )

( )

+ [

] *

( )

( )

( )

+) *

+

(2)

The Earth magnetic model in the simulator, however, is not

yet implemented. Therefore, the power consumption of

magneto-torquer operations are not simulated in real dynamics

of reducing angular momentum, but by assuming daily torque

operation maintenance using semi empiric model.

III. SIMULATION RESULTS

A. Zero Angular Momentum Mode

The zero angular momentum scenario only has 1 attitude

control mode, that is nadir pointing or 0 deg/s rate at Z & X

axis, and -0,06 deg/s at Y axis. The simulation starts with

assuming that the satellite already establish its X axis in flight direction and Z axis in nadir. Angular momentum of 0,1375

Nms with nearly equal vector component in each axis (X, Y, Z)

is then generated for the satellite, and the satellite is

commanded to maintain its nadir pointing. The reaction wheels

absorbed the angular momentum by rotating as commanded by

its controller, that has its feedback from the gyro at each axis.

For configuration 2D, simulation show that at steady state the

rotation of reaction wheel X, Y, and Z are -1000 rpm, -1100

rpm, and -944 rpm. Meanwhile for configuration 5D, they are

-985 rpm, -1085 rpm and -938 rpm. Table 2 shows the power

consumption of the reaction wheel (RW) per orbit for such attitude maneuver (coil = magneto-torquer).

The satellite is assumed to fly in low Earth polar orbit

(altitude of 650 km). It is assumed that the residual magnetic

dipole in the satellite interacts with Earth magnetic field and

increase the angular momentum. Commonly, in zero angular

momentum mode, gravity gradient torque also add to the

momentum. However, in this research, the effect from gravity

gradient is neglected. To reduce the angular momentum back to

its initial value, the 2 magneto-torquers are activated with total

power consumption of 5 Watts for 3 hours, every 2 days or 28

orbits. Calculated per-orbit, the additional power consumption is 0,5 Whr. The attitude sensor used is assumed to be gyros, star

Int'l Journal of Advances in Mechanical & Automobile Engg. (IJAMAE) Vol. 3, Issue 1(2016) ISSN 2349-1485 EISSN 2349-1493

http://dx.doi.org/10.15242/IJAMAE.IAE0716207 107

sensor, and horizon sensor (IR camera type), which per

reference [13] the power consumption is 18 Whr/orbit.

In table 2, the power consumptions from deployable panel

are compared with those of configuration with body mounted

solar panel from [4]. The table shows that the deployable

panels’ configurations have less reaction wheels power consumption than those of body mounted configurations,

which is due to lack of mode changing power consumptions

(hibernate to imaging) in the deployable panels configurations.

However, the deployable panels configurations have higher

total attitude control power consumption, due to longer use of

attitude sensors and computers.

TABLE II POWER CONSUMPTION PER ORBIT

Config RW coil total ACS total sat. (Whr) duty cycle

2 9 0 18 115 30%

5 9 0 18 115 30%

2D 5 0,5 23,5 205 60%

5D 5 0,5 23,5 205 60%

Table 1 shows that power production capacity of

deployable panels configurations are sufficient for operation of

60% payload duty cycle as noted in table 2. This mean that

configuration 2D could produce 5 million km2 of image with 13 m resolution, and configuration 5D could produce 2 million

km2 of image with 7 m resolution.

B. Angular Momentum Bias Mode

The angular momentum bias mode implemented in this

research has 2 attitude control modes, i.e., imaging mode and

hibernation mode. Since the imaging mode is done at 60% duty cycle, the satellite will be in nadir pointing attitude for 60

minutes or during entire sunlight. Figure 4 shows the Y reaction

wheel speeds during mode chaning operation.

Fig. 4. Reaction wheel speed during switching from hibernation to imaging mode for 5D configuration

Since the angular momentum bias mode is an open loop

system (from attitude sensors), the rigid body phenomena of eq.

(1) and (2) can clearly be seen. Fig. 5 shows that the satellite

nutates in X & Z axis due to cross product inertia components.

It is found that nutation rate is less than 0,11 deg/s at both the

hibernation and imaging mode. Such nutation, however, will

not degenerate the image quality, since payload operation

design has accommodate 20% pixel overlap, as noted in [14].

The figure also shows that the nutation frequency increase

during imaging mode

Fig. 5. Satellite angular rate during switching from hibernation to

imaging mode for 5D configuration

Table 3 shows the per orbit power consumptions from

angular momentum bias strategy. It shows that during

hibernation mode, the steady state power consumption of Y axis reaction wheels (RW) is 5,33 W. The Y axis reaction wheel

power consumption during imaging mode increase due to the

increase in the wheel rotation speed from 4000 to 5200 rpm (see

Fig. 4), to absorb the excess satellite rotation rate. Based on the

existing attitude control algorithm in the simulator, the transient

time for the mode switching is about 15 seconds.

The RW imaging power consumption in table 3 includes

transient power consumption during mode change from

hibernation to imaging (and vice versa as entering eclipse). The

power consumption from attitude control sensors/computers is

assumed to be the same as those in zero angular momentum mode. Also, it is assumed that the power consumption from

daily operation of magneto-torquer, that control the angular

momentum vector drift is 0,5 Whr/orbit.

TABLE III

PER-ORBIT POWER CONSUMPTION FROM MOMENTUM BIAS STRATEGY (60%

PAYLOAD DUTY CYCLE)

Config RW

hibernate

RW

imaging

coil total ACS

(Whr)

total sat.

(Whr)

2D 3,33 3,54 0,5 25,37 206,87

5D 3,33 3,54 0,5 25,37 206,87

Based on the simulations performed, the zero angular

momentum mode has almost the same power consumption than

the angular momentum bias mode. However, the momentum

bias mode has 2 inherent weaknesses, i.e. waviness in ground

scanning, which potentially reduced the image swath, and less

agile to take image at slightly off-nadir (needs more power for

slew maneuver due to the momentum stiffness). From system

reliability point of view, the zero angular momentum might has

1 weakness, i.e. depend non-interrupted attitude sensors

operation, which for micro-satellite is very hard to achieve. In

momentum bias mode, the necessary sun pointing will not be lost when the attitude control is interrupted.

time (s)

hibernation

imaging

time (s)

Int'l Journal of Advances in Mechanical & Automobile Engg. (IJAMAE) Vol. 3, Issue 1(2016) ISSN 2349-1485 EISSN 2349-1493

http://dx.doi.org/10.15242/IJAMAE.IAE0716207 108

IV. CONCLUSIONS AND FURTHER RESEARCH

The simulation shows that deploying solar panels in the Earth observation microsatellite configuration defined in this

research, can significantly increase the satellite performance. In

the design case, both configurations (2D and 5D) could achieve

twice the coverage compared to those of body mounted

configurations (2 and 5).

The research shows that LAPAN-ITB satellite simulator

can be used to further optimized Earth observation satellite

design by taking into account variables such as deployable solar

panels and attitude control modes. The case of placing the panel

on Y side, so it could face the Sun during morning pass while

maintaining nadir pointing, and the minimum numbers of

attitude control modes, have been proven to be efficient combinations, i.e. all power capacity increase from deploying

solar panel can be used fully for payload operation.

The trade-off parameters in between the 2 attitude control

modes are :

The zero angular momentum mode may lost its sun

pointing when the satellite’s attitude sensors the are

temporary interrupted, while the angular momentum

bias will not

The angular momentum bias mode has ground

scanning oscillation, due to its nutation, while the zero

angular momentum is not. The next research to be done is implementing the Earth

magnetic model in the satellite dynamic simulator, so that the

attitude maneuver using the magneto-torquer can be accurately

modeled.

ACKNOWLEDGMENT

The author also wish to thank Mr. Satriya Utama of LAPAN

for his advice regarding LAPAN-ITB satellite simulator

modifications.

REFERENCES

[1] W. Wu, H. Huang, S. Chen, B. Wu; ―Satellite Multidisciplinary Design

Optimization with a High-Fidelity Model‖; Journal of Spacecraft and

Rockets, Vol. 5, No. 2, 2013

http://dx.doi.org/10.2514/1.a32309

[2] J. T. Hwang, D. Y. Lee, J. W. Cutler, J. R. Martins; ―Large-Scale

Multidisciplinary Optimization of a Small Satellites Design and

Operation‖; Journal of Spacecraft and Rockets, 2014

http://dx.doi.org/10.2514/1.A32751

[3] R.H. Triharjanto; R.E. Poetro; S. Hardhientana, ―Multi-objectives

Optimization of Earth Observation Micro-Satellite Design Using Particle

Swarm‖, Proceedings of IEEE International Conference Aerospace

Remote Sensing, Jogjakarta, Indonesia, 2014

http://dx.doi.org/10.1109/icares.2014.7024371

[4] R.H. Triharjanto; R.E. Poetro; H. Muhammad, S. Hardhientana, ―Earth

Observation Micro-satellite Design Optimization Using Satellite

Simulator‖, Submitted to Journal of Mechanical Engineering (expected to

be published in 2016)

[5] E. Lorenz, H. Jahn, ―BIRD: More Than 3 Years Experience in Orbit,‖

Proceedings of the 5th IAA Symposium on Small Satellites for Earth

Observation, Berlin Germany, April 2005

http://dx.doi.org/10.1515/9783110919806.102

[6] H: S. Kim, T. J. Chung, N. H. Sung, K. H. Lee, ―Design Concept for

Autonomous Operation of KITSAT-3, an Experimental LEO

Microsatellite,‖ Proceedings of the IEEE Aerospace Conference, Big

Sky, USA, 2000

[7] H. Saito, N. Iwakiri , A. Tomiki, T. Mizuno, H. Watanabe, T. Fukami, O.

Shigeta, H. Nunomura, Y. Kanda , K. Kojima, T. Shinke, T. Kumazawa,

"High-Speed Downlink Communications with Hundreds Mbps from

50kg Class Small Satellites," Proceedings of the 63rd IAC (International

Astronautical Congress), Naples, Italy, 2012

[8] J. Eilers, J. Letschnik, U. Walter, H. Wusten, ―Analysis of the Power

System from the Earth Observation Satellite BIRD,‖ Proceedings of the

6th IAA Symposium on Small Satellites for Earth Observation, Berlin,

Germany, 2007

[9] B.J. Kim, H. Lee, S.D. Choi; ―Three-axis Reaction Wheel Attitude

Control System for KITSAT-3 Microsatellite‖; Space Technology 16,

vol. 5-6, pp. 291-296; 1996

[10] R.H. Triharjanto, M.S. Nur Ubay, S. Utama, R.E. Poetro, , ―Evaluation of

Attitude Dynamic Module on LAPAN-ITB Micro-Satellite Simulator‖,

Proceeding of IEEE International Conference of Aerospace & Remote

Sensing, Bali, Indonesia; 2015

http://dx.doi.org/10.1109/icares.2015.7429815

[11] M. A. Saifudin, M. Mukhayadi; ―LAPAN-A2 Attitude Control Strategy

for Equatorial Surveillance Mission‖; Proceedings of the 9th IAA

Symposium on Small Satellites for Earth Observation, Berlin, Germany,

2013

[12] M. Mukhayadi, R. Madina, U. Renner; ―Attitude Control of Bias

Momentum MicroSatellite Using Magnetic and Gravity Gradient

Torque‖, Proceeding of IEEE International Conference of Aerospace &

Remote Sensing, Jogjakarta, Indonesia, 2014

[13] S., Hardhienata, R.H. Triharjanto; ―LAPAN-TUBSAT System Budget,‖

in LAPAN-TUBSAT : First Indonesian Microsatellite, LAPAN 2007, ch.

4, pp. 26-32.

[14] R.H. Triharjanto; R.E. Poetro; H. Muhammad; ―Analisa Sensitivitas pada

Desain Awal Satelit Mikro Pengamat Bumi (Sensitivity Analysis on

Preliminary Design of Earth Observation Microsatellite)‖; Jurnal

Teknologi Dirgantara; Juni 2013.

Int'l Journal of Advances in Mechanical & Automobile Engg. (IJAMAE) Vol. 3, Issue 1(2016) ISSN 2349-1485 EISSN 2349-1493

http://dx.doi.org/10.15242/IJAMAE.IAE0716207 109