4 Final Report - Emits - ESA

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4 Final Report (Spacecraft Computer Unit) and instrument and high-speed interface for the instrument data to be transferred to the PDH. Cross-coupling is achieved by redundant Milbuses and cross-coupled interface to the PDH. Since G-Link is now obsolete, newer technology such as Aeroflex UT54 series may be more appropriate. The proposed electrical architecture is given in the following figure. The payload subsystem contains an Instrument Control Unit (ICU) for self-standing instrument mode control and to facilitate instrument testability. Figure 4.5-4 Geo-Oculus Electrical Architecture (Overview) 4.5.3 Power Subsystem The Electrical Power System (EPS) shall serve the satellite in sun and eclipse with the required power as derived from the power budget. The following essential EPS sizing requirements apply: • Life time: 10 years • Orbit: Geostationary • Maximum eclipse duration: 72min • Unit margin: between 5% to 30% pending maturity • System margin: 10% at EOL plus 10% time margin on energy recharge The main functions of the electrical power systems are • generation of electrical power with a photo-voltaic solar array • storage of excess energy during sun phases into the battery • safe distribution of the electrical power to the on-board users • protection of the battery against overcharging and deep discharging • fully autonomous operation. The EPS consists of solar array, battery, Power Shunt Regulator (PSR), battery charge and discharge regulators as well as bus distribution and protection. Two one axis drive mechanism provide optimum Page 4-52 Doc. No: GOC-ASG-RP-002 Issue: 2 Astrium GmbH Date: 13.05.2009
4 Final Report sun orientation for each of the two wings of the solar array. Power Budget Solar array and battery sizing assumptions: • Solar array temperature: 57°C • Summer solstice EOL • Solar array degradation for required life time • Unit power figures include maturity margins and an overall system margin of 10%. Power sizing The instrument power value is based on a mean figure without sun avoidance and a dual wing solar array configuration. For the other instrument/satellite alternatives (dual wing solar array with sun avoidance or single wing with/without sun avoidance) the solar array and the battery will be slightly smaller. This gives the following results: • Average load power: 1800W • Required solar array area: Minimum 10.1 m 2 • Battery size: 135Ah (11 string 3 parallel configuration) • Battery mass: 48kg W 2500 2000 1500 1000 500 Figure 4.5-5: Power Profile 0 Power Profile GEO-Oculus 396 796 1196 min 100,00 90,00 80,00 70,00 60,00 50,00 40,00 30,00 20,00 10,00 0,00 Power SA W Power profile Load W Battery SoC The solar array area includes 3.5% margin for string failures. The battery is based on G5 technology. Cell losses are covered by two additional modules (parallel cells). The maximum battery DoD is approx. 62%. The battery sizing is based on nominal in-orbit operations. 4.5.4 Payload Data Handling and Transmission Data rate assessment For the sizing of the PDHT, an average data rate of 250Mbps coming from the instrument is assumed. Doc. No: GOC-ASG-RP-002 Page 4-53 Issue: 2 Date: 13.05.2009 Astrium GmbH SoC
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Geo-Oculus: A Mission for Real-Time
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DL Final Distribution List Report Q
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TOC Final Table of Content Report D
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1 Final Report 1 Introduction 1.1 S
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1 Final Report [RD 7] Candidate Mis
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4 Final Report - Emits - ESA

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