Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1): 26- 33
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Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
Development of Software-Based Intelligent Power Storage and
Management Control System (IPSMCS) For Satellite Application
Adegboye, K.A. And Baruwa, A. A.
Department of Electrical-Electronics Engineering
Osun State College of Technology,
P.M.B.1011, Esa-Oke, Osun State, Nigeria.
Corresponding Author: Adegboye, K.A.
_________________________________________________________________________________________
Abstract
It is imperative that electrical power supply reliability and stability is pre-eminence for satellites to navigate
their orbits and achieve their various application objectives. The Satellite Electrical Power Subsystem (EPS)
generates, stores, controls, conditions and effectively distributes power for satellite’s housekeeping operations.
The motive of this research work is to develop a software-based intelligent Power Storage and Management
Control System (IPSMCS) for application in the intelligent coordination of power storage, power savings, load
switching and control of battery temperature in the EPS of a satellite. However, many application programs are
suitable in this application such as; Fortran programming language, C-programming language, Assembly
programming language and among others. But in this work, C-programming language was considered to
develop the IPSMCS program codes to be implemented in KT69A53 microcontroller for software-based
intelligence while Java programming language was used to achieve a simulation model for virtual test for
system efficiency. The IPSMCS managed power through load shedding and power-cycling and actuated fault
protection mechanism during power emergencies very quickly and automatically especially during a natural
phenomenon dark period such as eclipse and sunlight orbit periods. It provides intelligence for real-time power
storage. It monitored the solar source outputs to determine orbit periods, so also is the battery charge and
discharge processes to determine threshold levels for Battery State Of Charge (SOC) and Depth Of Discharge
(DOD) and actuated suitable battery charge rate regime for effective system performance. The IPSMCS offered
an improvement over the conventional satellite power management and control systems with regards to the
system’s size reduction, self-contained intelligence, flexibility and improved system response to power
emergencies. In this regard, a sustainable reliability of orbit operation is guaranteed and assured.
__________________________________________________________________________________________
Keywords: spacecraft, satellite, electrical power supply, orbit, software, battery, solar array.
outer space. The ill-fated Nigerian Communication
Satellite (NigCom Sat) launched by China which got
missing in operation after two years was reported to
have suffered a failure due to power related technical
hitches as reported by (Nkanga; 2008).
This
unsavoury development is a proof that in the event of
system and environmental emergencies in the outerspace, the Satellite’s EPS is most vulnerable.
Therefore, spacecraft power system must be
incorporated with some levels of intelligence that will
enable the satellite to sense undesirable systems and
environmental changes more quickly and accurately
as well as take appropriate corrective actions in order
to make it durable and usable for its lifespan without
destruction which is sometimes of alarming
dimension (Pratt and Bostian, 1986).
INTRODUCTION
The term “Satellite” is typically used to describe a
spacecraft invented on the Earth and sent into orbit
on a launch vehicle to perform some specific
missions or tasks.
Typical examples of these
spacecrafts are navigation satellites, weather
satellites, communication satellites and scientific
satellites. Technically speaking! Anything that is in
orbit around the Earth is a Satellite. The need for
reliable electrical power supply is very primary for
satellites to navigate their orbits and achieve their
desired objectives in accordance with the claim of
Okoro, Okafor and Ejimanya (2007). The Satellite
Electrical power subsystem (EPS) is responsible for
electrical power generation, energy storage for peak –
power demands and eclipse periods, power regulation
to prevent overcharging and undesired spacecraft
heating, power switching and distribution to other
systems and subsystem (Corbin Cote and Ulrich;
2006).
Consequently, power failure which may be resulted
in spacecraft accident can be averted since apart from
weather or atmosphere conditional problem, power
failure has been accounted for the gross spacecraft
accidents of the recent time.
Satellite systems in orbit are often subjected to the
hostile hard vacuum environmental conditions of the
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Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
-
OVERVIEW
OF
CONVENTIONAL
SATELLITE EPS CONTROL
In the conventional satellite EPS, the system is
commonly designed and configured as indicated in
the highlights below:
- Power Management embedded software program
is situated in the spacecraft on-board Central
Processing Unit (CPU) by Bonnema; (2003). As
a result, system intelligence is not domiciled in
the satellite EPS. Therefore, any CPU related
problem is immediately translated to the EPS
- EPS system actuators are controlled by hardware
logic gates circuit and relay switches (Panneton
and Jekins; 1996). This resulted in limited
accuracy and speed of system response to power
emergencies. Another problem here is an
exceeding system site and weight. Whereas, size
and weight reduction is an essential design
consideration in satellite development.
To achieve system’s size and weight reduction.
EVOLVING TECHNOLOGY
In the modern world of satellite system development,
the concept of “simulating before developing” is of
great economic importance. It is therefore crucial to
have all of the intended system design parameters
simulated and optimized before a heavy commitment
to implementation. In the light of this, the research
covers as follows:
- The development and implementation of
software-based intelligent power management
control system using C codes embedded program
routine for 8-bit microcontroller.
- The development of 8-bit microcontroller
hardware circuit structure upon which the
software
system
development
and
implementation is based.
- The Java program simulation of the control
system and environmental variables of the outer
space. The simulation model is for virtual testing
of the functions and effectiveness of the IPSMCS
as the real space test is difficult and almost
impossible.
AIM AND OBJECTIVE
The pre-occupation of this paper is to evolve software
based Intelligent Power Storage and Management
Control
System
(IPSMCS)
for
automatic
coordination of power storage (battery charging),
power savings, and load switching and
voltage/temperature control functions of the EPS
using C-program routine. However, the fundamental
objectives are as follows:
- As a better replacement to hardware logic gates
circuits located within the EPS with C-program
code embedded for improved system response to
power emergencies.
- To achieve a self contained and autonomous
system independent of the satellite on board
Central Processing Unit. (CPU).
- To isolate CPU related problems from adversely
affecting the performance of the EPS.
SATELLITE SUBSYSTEM OVERVIEW
All the sub-systems of all satellites have need for
electrical power to sustain their operations while in
space. A satellite is divided into two parts.
(i)
The payload part
(ii)
The bus part
The payload part principally consists of antennas and
transponders in communication satellite studies while
bus part essentially consists of six sub-systems as
shown in figure 1.0 below as depicted from the work
of Takashi (2006).
Antennas
Payload Part
Transponders
Attitude Control System
Artificial Satellite
Propulsion System
Power Supply System
Bus Part
Thermal Control System
Structural System
Fig. 1.0: Structure of a Communication Satellite’s Sub-System
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Telemetry/Command System
Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
terrestrial station can issue commands to the
satellite for operational control, such as turning
parts of the equipment ON and OFF, and
telemetry functions for transmitting data on the
status of the satellite, such as equipment
temperature, fault condition and power
management and distribution report.
PAYLOAD PART
The payload part comprises of the equipment
responsible for achieving the satellite mission
objectives such as scientific, navigation payloads,
communication, etc.
BUS PART
The components of the bus part are as described
briefly as follows:
- The Attitude Control system is used to keep the
satellite pointing in right direction and orientation
by overcoming attitude disturbances such as
pressure of solar radiations and gravitational
gradients.
- The propulsion system placed the satellite in the
prescribed orbit and control its orbit with respect
to orbital perturbations arising from the sun and
moon and from the flatness of the earth.
- The Electrical Power System supplies power to
the satellite. This power is obtained from solar
cells in sunlight, and from arrays of batteries
when shaded from the sun.
- The thermal control system keeps the temperature
of all equipment on-board the satellite within a set
temperature range safe for the equipment.
- The structural system maintains the overall
integrity of satellite throughout the vibration and
acceleration it is subjected to during launch.
- The telemetry, tracking and command (TTC)
system has a command functions whereby a
SATELLITE ELECTRICAL POWR SYSTEM
(EPS) IMPLEMENTATION
The operation of the satellite components are
energized by the EPS. Therefore, the EPS must be
developed to function optimally and reliably
throughout the period of satellite mission. The
satellite Electrical power subsystem is responsible for
electrical power generation, energy storage for peakpower demands and eclipse periods, power regulation
to prevent overcharging and undesired space craft
heating, power-switching and distribution to other
sub-system as well as power management (Ulrich,
Corbin and Cote; 2006). To achieve all these
assertions of thought, the EPS incorporate both solar
panels for power supply while the satellite is in
sunlight and batteries for storing and providing power
when the space craft may be in Earth’s shadow. The
EPS functions requirement as indicated by Ulrich,
Corbin and Cote; (2006); is as shown in figure 1.1.
below.
Supply power to spacecraft bus and payload
Generate power
Store power
Regulate power
Distribute power
Regulate the power generated
by the power source
Regulate the bus voltage
Charge the battery
Fig. 1.1 EPS functions requirements
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Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
EPS METHODOLOGY
There are five major space power generation sources;
solar power, fuel cells, batteries, nuclear and
microwave. The choice of an appropriate power
generation system depends on the amount of power
required. Renewable energy source from the sun such
as the solar arrays distinguishes itself as the energy
source of preference in the solar space by Harty,
Otting and Kudija; (1993),
vii)
It is however evident that the issue of electric power
generation, storage and distribution in satellite
demands intelligent managements and control. The
essence of the intelligent power management and
control system is to ensure the following:
Monitoring of the entire EPS performance in
relation to the desire objective; battery
amount and voltage charge / discharge ratio,
threshold values and limits.
Runs embedded decision program and issue
control instructions.
Power stability, reliability and adequate
system protection at all time.
Fault isolation and system recovery and
power regulation and economics in
compliance with specification and standards.
viii)
ix)
x)
xi)
EPS COMPONENTS DEVELOPMENT
OVERVIEW
i)
Solar Array (SA): This is an array of PV
cells that convert sunlight into electricity.
ii)
Solar Array Drive: The solar array drive
(SAD) consists of slip rings, a motor, and
motor drive electronics. It orients the solar
array to face the sun for generating the
maximum possible power during the entire
sunlight period of the orbit.
iii)
Shunt
Dissipator:
During
sunlight,
particularly in the beginning of life, this
compound dissipates power that is unwanted
after meeting the load power and the battery
charge power requirements.
iv)
Battery: The battery stores energy in an
electro-mechanical form to supply power to
the loads during eclipse periods over the
entire mission life.
v)
Power Regulation Unit: The power
regulation unit (PRU) provides an interface
between the solar array bus and the battery.
The battery voltage varies widely with the
cell voltage varying from 1.0V when fully
discharged to 1.55V when fully charged.
The discharged converter in the PRU boost
the battery voltage to the bus voltage during
an eclipse and the charges converter boost
the array voltage to the battery voltage
during sunlight.
vi)
Power Distribution Unit: The power
distribution unit (PDU) ensures that all
loads, except critical and essential loads are
xii)
xiii)
powered through switches and fuses. The
fuses are to protect the power system from
faults in the user equipment more than
protecting the loads.
Bus Voltage Controller: This consists of the
bus voltage sensor, the reference voltage and
the error signal amplifier. The amplifier
error signal output of the bus controller
enters the mode controller, which in turns
sends command signals needed to regulate
the bus voltage within required limits.
Mode Controller: The mode controller
automatically changes the EPS mode in
response to the error signal with the aid of
MOSFET switch as considered in this
research for better performance.
Battery Bus: A tap point directly off the
battery. During the launch and ascent
phases of the mission, the PV array is not
deployed, and the battery meets all the
energy needs through the bus.
PRU Bypass Mode: The PRU control
quickly delivers the battery energy to the
fuse in case of a fault in any of the loads to
blow off as quickly as possible in order to
minimize the bus voltage decay.
Shunt Mode: During sunlight, if the solar
power exceeds the load and battery charge
requirements, the mode controller turns on
the shunts to dissipate excess power,
otherwise, the bus voltage will rise above
the allowable limit. At this mode, the
battery is charged at full, reduced (cut back)
or trickle charge rate.
Charge Cut-back Mode: When the battery
is approaching full charge, the charge rate is
cut back to control the battery temperature.
Discharge Mode: In the absence of solar
power during an eclipse, the battery is
discharged to maintain the bus voltage.
POWER STORAGE AND MANAGEMENT
SOFTWARE
Power and Energy Management Software (PEMS) is
part of the software system dedicated to the
performance of EPS, health monitoring, control and
protection. In case of emergencies, or during planned
mission operations, PEMS sheds loads in a preset
sequence, when the battery state of charge cannot
support all loads. The battery telemetry consists of
the battery voltage, current, temperature, individual
cell voltages and the internal pressures of the selected
cells. Most of these telemetry readings goes to
PEMS in accordance with Patel; (2005).
LOADS
The term load includes all loads as; the payloads
(receivers, transmitters, measuring instruments, etc)
and the bus system loads. Most loads in satellites are
constant power loads. The orbit average power
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Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
requirement is taken into account in sizing PV array
and the battery for all loads.
•
GROUND POWER CORD
In order that, the battery power is preserved during
pre-lunch testing and final checks before lift-off, the
on-board system uses external ground power via an
umbilical cord. To further preserve the battery,
power transfer is scheduled as late as possible in the
countdown.
•
BUS VOLTAGE CONTROL
As earlier mentioned, shunt mode regulates the bus
voltage by using a shunt dissipator which is the tool
employed in this paper.
CONTROL CIRCUIT
The control circuit employed in this power is digital
control circuit.
Digital control circuit is an
alternative to the traditional analog control. A digital
shunt regulator uses relatively small shunt dissipator
to provide small signal control of the bus. Analog
shunt is unstable in response but offers much better
dynamic response to the power bus transients. But
digital control is preferred in this research as it offers
flexibility in tailoring the system to multiple
missions. By simply changing the gain constant in
the software table, one can adjust the system’s
transient response. One can also incorporate a
number of different battery charge regimes, and then
adjust the charge rate in orbit with single command
from the ground. This arrangement is adaptable to a
variable number of solar array segments and to a
number of array configurations. The result is more
flexibility in using standard modules of the solar
array and the battery, which translates into cost
reduction (Patel; 2005).
•
•
PSMCS MODULAR FUNCTIONS IN
SUMMARY
The modular functions of the PSMCS are categories
as:
(i)
Battery Charge/Discharge
(ii)
Power Monitoring
(iii)
Power Savings
(iv)
Voltage/Temperature Control
(v)
Orbit Period Control
The functions which can however be summarized as
shown in figure 1.2 and elucidated thus:
•
Battery Charging/Discharging Module: The
battery charging module is designed to allow
the EPS battery to be intelligently and
effectively charged and maximized their
lifetime. All functions necessary to charge the
30
batteries intelligently are implemented in a
battery charger chip and 8-bit microcontroller.
Power Monitoring: The Power System may
literally be tasked beyond what is capable of
generating when all systems are in operation.
Therefore, several procedures are implemented
to allow the power system to function
adequately and the mission to succeed.
Power Saving Module: Shunt module is
employed for this purpose as discussed earlier
on. Moreover, during eclipse periods and
there is a significant drops in battery below a
tolerable limit indicating a low “state of
charge” (SOC), the PSMCS will respond by
shedding non-critical loads in idle mode. This
reduces or sheds the load on the battery and
allows its voltage to recover from the
undesirable low level.
Voltage/Temperature Control: In most
cases, the power generated by solar array
exceeds the power requirement of the
subsystem loads that excess is taken up by the
batteries. The purpose of the PSMCS focuses
in this paper is to manage the dissipation of
excess solar array power not needed for battery
charging. This is achieved by software control
f digital switches to shunt out individual solar
array circuits to regulate supply of power for
battery charging. To best maintain battery
health, the PSMCS provides automatic charge
cutback if it detects excessive battery voltage,
pressure, or temperature.
Orbit Period Control: The Satellite orbit
induces thermal cycling of the spacecrafts
components. The solar array and other parts
that extend from the spacecraft’s body and
have comparatively low heat capacity are
subjected to temperature extremes larger than
those experienced by the rest of the spacecraft.
When the spacecraft exits the Earth’s shadow,
the solar panels temperature may be as low as 80oc. But on re-entry into sunlight they
rapidly according to the research finding,
peaking in the worst case at approximately
65oc according to Panneton and Jekins; (1996).
The satellite mode of operation induces
thermal concerns from within the spacecraft
and this EPS thermal concern is the major
concern of this research as tremendous
temperature or internally produced heat affects
component performance and any elevated heat
generation must be removed as quickly as
possible from the spacecraft.
Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
Fig. 1.3: IPSMCS in charge mode during sunlight period
Battery Discharge Mode: During eclipse period, the IPSMCS detected this condition by monitoring the output
of the solar source through a current sensor and actuated the battery discharge mode through a microcontroller
energized digital switch to supply power to the satellite as shown in figure 1.4.
Fig. 1.4: IPSMCS in discharge mode during eclipse period
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Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
Solar Output Monitoring: In the event of excessive
solar power supply, IPSMCS through the current
sensor connected to the output of the solar source
detected this condition and energized a shunt circuit
to absorb excess power to prevent it from destroying
the system. On the other hand, in the event of
insufficient solar power supply, IPSMCS actuated
battery discharge mode to make up for the
insufficient power supply to satellite loads for
reliable and efficient performance, this is shown in
figure 1.5.
Fig. 1.5: IPSMCS in Shunt mode in the event of excessive solar supply
LIMITATION
To avoid software pulled interrupt, the application
packages should be saved on electronic-erasable part
of memory (Eeprom) and the battery for the system
must be of high quality with appropriate rating.
REFERENCES
Bonnema, A. R. (2003): The Delfi – C3 Student
Nanosattelite; IAC – 05, B5, 6A, 02.
Harty, R. B., Otting, W.D., and Kudija, C.T. (1993).
“Application of Brayton Cycle Technology to Space
Power” Proceedings of the 28th Intersociety Energy
Conversion Engineering Conference, Atlanta, Aug. 8
– 13, 1993, Vol.1.
CONCLUSION AND RECOMMENDATIONS
The paper achieved the Software-Based Intelligent
Power Storage and Management Control System
(IPSMCS) which is an improvement over the
conventional satellite power management and control
system for application in the automatic coordination
of real time power storage, power savings, load
switching and battery temperature control in a
satellite EPS. IPMCS also monitored the solar source
output to determine the orbit period. It monitored
battery charge and discharge processes to determine
threshold levels or Battery State of Charge (SOC) and
Depth of Discharge (DOD) and actuated suitable
battery charge rate regime for reliable system
performance. Battery temperature threshold level is
also controlled to control destructive exothermic
reaction. It can however be recommended that, the
research has created a ground for further research in
the implementation of the KT69A53 microcomputer
hardware technology and further study is desirable in
desirable in EPS thermal concerns of spacecraft.
Nkanga, E. (2008): “Nigeria N40billion Satellite
Missing from Orbit”.
Okoro, C.E., Okafor, E.N, Ejimaya, J.I, (2007):
“Design Algorithm of Electrical Power Generation,
Distribution and Control for Satellite System”. 1st
Proceedings of National Centre for Satellite
Technology Development, Abuja.
Panneton, P.E. and Jekins, J.E. (1996): “The MSX
Spacecraft Power Subsystem” John Hopkins APL
Technical Digest, Volume 17, Number 1; pp.78.
Patel, M.R. (2005): “Spacecraft Power System” CRC
Press.
Pratt, T. and Bostian, C.W. (1986) “Satellite
Communications”
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Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1):26- 33 (ISSN: 2141-7016)
Tashi, I. (2006): “Satellite Communications Systems
and its Design Technology.
Ulrich, S., Corbin, F.L, Cote, H.N. (2006):
“Conceptual Design of Electrical Power System for
the European Student; Moon Orbital Mission” 57th
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C3.1. 07, Spain.
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