Journal of Emerging Trends in Engineering and Applied Sciences (JETEAS) 6(1): 26- 33 © Scholarlink Research Institute Journals, 2015 (ISSN: 2141-7016) jeteas.scholarlinkresearch.com 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 26 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 27 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 28 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 29 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 31 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” 32 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 International Aeronautical Congress, IAC – 06 – C3.1. 07, Spain. 33
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