imaging satellite mission: eyeon

EYEON GMBH Proposal for a highresolution multi-temporal imaging satellite mission: eyeon ESPACE Satellite Mission Design Project Meltem Eren, Fathalrahman Adam, Sean Hannon, Denise Schmidt, Nico Trebbin 29.06.2012

Executive Summary Earth observation satellites have emerged as the preeminent tool for consistent and high-resolution imagery. The eyeon mission is designed to provide rapid on-demand sub three meter resolution imagery as well as generate a database of regions of interest. In the following document, the eyeon team proposes a multi-use satellite observation system capable of servicing many constituents in different areas of study. The consistent illumination conditions experienced by the satellites can provide similar illumination conditions across the mission lifetime. Customers requesting images of crops may analyze growth cycles at monthly and yearly intervals. The land usage in developing regions is observable as well. This type of data in the visual and near infrared spectrum is of particular interest to land management programs of governmental agencies and agricultural customers. The constellation of five satellites in a walker orbit enables the entire Earth to experience an overflight within a 24-hour period. This short revisit rate carries distinct advantages over other systems when the satellites are tasked with capturing images of disaster-affected areas. The quick response time of one day enables the distribution of level 3 image products to the customer within one day of request to image. A multispectral camera on each satellite captures images at a high resolution useful to insurance agencies in assessing damages. Objects such as damaged buildings or burned ground areas are visible in the eyeon products. Combination of the images taken from the eyeon system with imagery in other spectral ranges can greatly enhance utility. The high-resolution visible range images of eyeon may be combined with SAR products to provide three dimensional surface models. As a full member of the International Charter on Space and Major Disasters, the DeutschesZentrumfürLuft- und Raumfahrt (DLR) would increase its contribution through the eyeon mission. As the use of satellite imagery grows, the DLR would remain a key player in the contribution of data to respond to the average of 300 disasters per year serviced by satellite missions. The combination of high optical and temporal resolution provides the eyeon mission with distinct advantages over existing missions. A steady demand for satellite imagery and a flexible platform Page 1

Table of Contents Executive Summary... 1 1. Objectives and Requirements... 4 1.1 Objectives... 4 1.2 Mission Requirements... 4 1.2.1 FunctionalRequirements... 4 1.2.2 Operational Requirements... 4 1.2.3 Constraints... 4 1.3 Background... 4 1.3.1 Existing Imaging Missions... 4 1.3.2 Characterization of Area of Interest... 6 1.4 Applications... 8 1.5 Retrieval Algorithms... 9 2. Mission Assumption and Technical Requirements... 10 2.1 Mission Characteristics... 10 2.2 Mission Simulations and Technical Requirements... 12 2.3 Complementary Data and Synergies... 12 3. Proposed Mission Architecture... 12 3.1 Space Segment Technical Concept... 12 3.1.1 Tradeoffs for spacecraft subsystems design... 12 3.1.2 High-resolution multi-spectral imaging system... 13 3.1.3 Flight Commander... 15 3.1.4 On-board communication... 15 3.1.5 Power subsystem... 16 3.1.6 Spacecraft power requirement calculation... 16 3.1.7 Power regulation... 18 3.1.8 Solar array design... 18 3.1.9 Battery system... 19 3.1.10 Power cycle calculation... 20 3.1.11 Thermal subsystem... 21 3.1.12 Basic Energy Balance... 22 3.1.13 Attitude Determination and Control System (ADCS)... 25 3.1.14 Structure... 29 Page 2

3.2 System Budgets... 30 3.2.1 Mass Budget... 30 3.2.2 Power Budget... 31 3.3 Ground Segment Elements... 32 3.3.1 Ground Stations... 32 3.3.2 Mission Control Center in Munich... 32 3.3.3 Ground Station in Svalbard/Norway... 33 3.3.4 Ground Segment Architecture... 33 3.3.5 Up/downlink... 34 3.3.6 Data Handling... 35 3.4 Mission Phases Summary... 35 3.5 Launcher... 35 4. Programmatic Elements... 36 4.1 Product Description and Delivery... 36 4.2 ROM Cost Estimate... 40 4.2.1 Ground Segment... 40 4.2.2 Subsystems... 40 References... 42 Annexes... 44 Page 3

1. Objectives and Requirements The following objectives and requirements were generated based on the ITT document distributed during the introduction of the ESPACE Satellite Mission Design Project. 1.1 Objectives 1. To capture images of the earth of interest to the agricultural community, insurance agencies, and disaster response/relief efforts. 2. To provide high and low level multispectral image data to customers as requested. 3. To service customers in research institutes, at governmental agencies and private enterprises. 4. To provide the image information in a timely manner relevant to the customer need. 1.2 Mission Requirements 1.2.1 FunctionalRequirements FR1 -The imaging system must generate images at 5m resolution. (Sufficient for disaster response) FR2 - The imaging system must be capable of capturing spectra necessary to provide customer requested data products. 1.2.2 Operational Requirements OR1 - The system must revisit any location in 24hrs. OR2 - The customer products must be delivered in 4 days. OR3 Lifetime of 5 years 1.2.3 Constraints C1 - Images may only be taken during daylight hours. C2 - Existing Technology must be used. C3 - Cost must be held to reasonable level. 1.3 Background 1.3.1 Existing Imaging Missions In order to get familiar with the topic of designing a satellite mission that is capable of taking high resolution images a background study on the already existing systems IKONOS, RapidEye and QUICKBIRD was conducted. The most important features are compared in table 1. IKONOS and QUICKBIRD are heavy single spacecrafts as opposed to RapidEye which is a constellation of five small satellites in one plane. All three missions have global coverage due to their high inclination and polar, sun-synchronous orbit configuration. RapidEye and QUICKBIRD are able to offer a daily revisit time which will also be crucial for the eyeon mission concept whereas IKONOS only has a spatial resolution of 3-5 days. The swath width is also different for each of these three missions, as Page 4

well as the possible spectral resolutions. IKONOS only has a swath width of 11 km but it can offer a resolution up to 0.82 m in the panchromatic band and 3.2 m in multispectral. RapidEye s swath width is the biggest (77 km) but the spectral resolution is only 6.5 m in the visible band. QUICKBIRD is a high resolution system with up to 0.61 m resolution for the panchromatic and 2.44 m for the visible band. The swath width is 16.5 km. All three systems have a multispectral (MS) camera which covers the visible (VIS) and near infrared (NIR) band; IKONOS and QUICKBIRD offer panchromatic (PAN) and RapidEye offers Red Edge in addition. Table 1: Comparison of important features for other imaging satellite missions IKONOS RAPIDEYE QUICKBIRD ORBIT altitude 681 km 630 km 450 km polar (i = 98.1 ) polar (i = 97.8 ) polar (i = 97.2 ) sun-synchronous sun-synchronous sun-synchronous revisit time 3-5 days off-nadir daily off-nadir 1-3.5 days off-nadir 144 days true-nadir 5.5 days true-nadir SPACE lifetime 5-7 years 7 years 7 years CRAFT mass 817 kg 5 x 150 kg 950 kg body size 1.8 x 1.8 x 1.5 m 1 x 1 x 1 m 3.04 m length launcher Athena-2- rocket Dneprt Boeing Delta II launchsite Vandenberg Air Force Base (California) Bikonur Vandenberg Air Force Base (California) COMMU NICA uplink S-band at 2 kbit/s (command) S-band at 9.6 kbit/s S-band at 2 kbit/s TIONS downlink X-band at 320 Mbit/s (imaging data) X-band at 80 Mbit/s (imaging data) X-band at 320 Mbit/s (payload data) S-band 32 kbit/s (housekeeping) S-band at 38.4 kbit/s X-band at 256 kbit/s (housekeeping) on board 48 Gbit, 1500 km of data storage 64 Gbit image data per orbit 128 Gbit dynamic range 11 bits per pixel 12 bits 11 bits SENSOR swath width 11 km 77 km 16.5 km accessible ground swath 700 km (off nadir at standard angles) spatial resolution 0.82m (nadir) - 1m (26 off-nadir) panchromatic 3.2m (nadir) - 4m (26 off-nadir) multispectral 6.5 m (nadir) 544 km (30 off nadir) 0.61m (nadir) - 0.72m (25 off-nadir) panchromatic 2.44m (nadir) - 2.88m (25 off-nadir) multispectral 1 m pan-sharpened spectral resolution PAN, VIS, NIR VIS, Red Edge, NIR PAN, VIS, NIR Page 5

1.3.2 Characterization of Area of Interest The eyeon mission is designed to deliver image data to customers in the field of agriculture, insurance and disaster management. To accomplish this goal different factors have to be taken into account. Insurance companies only need information about inhabited areas. Therefore, figure 1 shows the Earth s population density. It is obvious that Greenland, the northern parts of North America and Asia as well as desert areas (e.g. Sahara) are of less interest. Areas with a big risk for natural hazards (figure 3) also need to be investigated. It is important for insurance company to know this risk and investigate further. For agricultural applications only vegetated areas are of interest. This mainly includes forest areas as well as cropland; barren and sparsely vegetated land is unlikely to undergo big changes, hence less important for agricultural studies. Figure 2 is a map of the global land cover distribution showing that also the northern areas of the globe and the deserts are no areas of interest. Disaster management is only necessary in inhabited areas (figure 1), but also depends on the occurrence of disasters. Figure 3 is a map of different types of disasters which happened in 2011. It shows that the most northern areas of the Earth are unlikely to be hit by a natural disaster. Figure 4 is also related to disaster management. It shows the major shipping routes and is thus related to the risk of having to detect an oil spill. To conclude, the main areas that need to be covered by the eyeon mission are between -70 and +70 latitude. That way, all customers in the field of agriculture, insurance and disaster management will be provided with accurate data. Figure 1: Global map of population density showing the persons/squkm (Source: FAO 2005) Page 6

Figure 2: Map of global land cover distribution by dominant land cover type (Source: FAO 2006) Figure 3: Global distribution of natural catastrophes in 2011 (Source: 2012 Münchener Rückversicherungs-Gesellschaft, Geo Risk Research, NatCatSERVICE) Page 7

Figure 4: Global shipping routes (Pablo Kaluza et al. 2010) 1.4 Applications The eyeon constellation is able to reach any area within ±70 latitude within 24 hours. The camera mounted on the spacecraft is able to cover the visible, panchromatic and near infrared band. The high temporal, spatial and spectral resolution allows eyeon imagery to be used in various fields of application. The use of multispectral imagery enables the customer to distinguishing between different wavelengths of electromagnetic radiation. With a maximum resolution of 1.5 m the eyeon system ensures to covers most of the applications in the fields of agriculture, insurance and disaster management. The only constraint is given by the used wavelengths. In the field on agriculture the limitations are small. Soil moisture is unlikely to be estimated accurately, imagery from a satellite using SAR sensor should rather be used for it. Limitations in the field of disaster management are for example hot spot detection or flood disaster management. Hot spots can only be detected in the IR range (Nirupama 2002). Water bodies are better detected by SAR sensors and the cloud coverage during most floods events blocks the view of the MS sensors. Volcanic ash is also unlikely detected; MIR sensors are usually used for it. However, lava flows can be identified in the NIR band (Nirupama 2002). Despite these limitations, eyeon imagery is the perfect choice for many applications. The choice of sunsynchronous orbits results in consistent illumination conditions which is an advantage for agricultural research. EyeON imagery enables direct observation of the land surface which can be used for crop assessment, crop health, change detection, forestry or monitoring of land use in developing regions. The eyeon satellites can revisit any location on Earth within 24 hours. This short response time is especially useful for disaster management. Images have to be delivered to the Page 8

customer within a short time period to do damage assessment and relief management. Possible disasters are floods, droughts, hurricanes, earthquakes, landslides, volcanic eruptions, fires or oil spills. Satellite imagery plays an important role in risk assessment as well. This is particularly interesting for insurance companies. Also the images can be used for preparedness and prevention purposes. In the post-disaster phase images can provide information on gathering sites or monitor the reconstruction measures. In combination with imagery from missions that record other spectral ranges, the number of applications of the eyeon imagery increases. For instance the combination with SAR images allows to calculate a three dimensional surface model or to precisely assess flooded areas. 1.5 Retrieval Algorithms Figure 5: Low level ground segment architecture The figure above provides an illustration of the planned data dissemination algorithm. The eyeon Ground Segment is designed in such a way, that we will download the acquired data stored on the satellite, at each passing of each single satellite to our Ground Station. With this design approach of the eyeon Ground Segment we can optimize the degree of capacity utilization for the dedicated Ground Station such that we have almost no idle time. So this approach is very cost-effective. For redundancy purposes the system is designed in a way, that we can also use a 3 rd party Ground Station Page 9

operated by ESA geographically located right next to the self-operated Ground Station. The operational control is done via the Mission Operation Center located in Munich. Therefore eyeon uses a WAN connection to the antennas, the mainframe and the dedicated physical hardware installed in Svaldbard. So, with this design approach of the Ground Segment we can ensure operation, even if the terrestrial link (WAN) fails and the connection to the Ground Station is lost. For this purpose eyeon is going to establish an emergency on-demand service with limited personnel resident to provide a 24/7 emergency plan. 2. Mission Assumption and Technical Requirements 2.1 Mission Characteristics In the design of the eyeon constellation many different configurations were considered. After estimating approximately five satellites to cover the region of interest, three constellations were developed. A five-satellite constellation in a polar street configuration was compared with a fivesatellite walker constellation. The walker constellation was ultimately selected based on greater visibility of the satellites from the Svalbard station. The 97.7 degree inclination of the individual walker constellation member additionally provided sun-synchronous movement a key advantage addressed later in detail. Figure 6: Polar View of eyeon Constellation (STK ) The selected orbit must fulfil the following requirements: 1. The constellation must enable global coverage in a 24-hour period. 2. The orbit altitude must accommodate the selected imaging system. 3. The constellation must enable regular and quality access to the ground station network. Page 10

eyeon The following table shows orbit characteristics for an individual eyeon satellite. Each satellite is spaced at 72 deg in the RAAN. Table 2: Orbit characteristics for each eyeon satellite Property Orbits per Day Repeat Cycle Orbits in Cycle Orbit Period (minutes) Inclination Semi-Major Axis [Orbit Radius] (km) Orbit Velocity (km/s) Mean Altitude (km) Value 14.86 31 461 96.6 97.7 6878 7.5579 600 The coverage pattern of the selected constellation fulfils the required revisit rate. The following image depicts the 100% coverage of the Lat. [70,-70] degree area of interest. Figure 7: Coverage Area in blue of eyeon Constellation in 24hr Epoch (STK) Figure 8 shows the position and velocity differences between perturbed and unperturbed orbit for 24 hours. Considered perturbations are atmospheric drag, solar pressure, direct tides of Sun and Moon, solid earth tides, pole tides and ocean tides under GRIM-C5 gravity model. The major impact in the perturbations is caused by atmospheric drag since the satellites will be on low earth orbit. Page 11

Figure 8: Difference in position and velocity of perturbed and unperturbed orbit 2.2 Mission Simulations and Technical Requirements In order to compare the constellation concepts and evaluate the selected mission orbit characteristics, a model was developed using the AGI STK 9 software suite. The STK program enables visualization of the satellite constellation, ground tracks, simulation of coverage, ground segment contact periods. MATLAB was also used to calculate perturbation effects and other general calculations. 2.3 Complementary Data and Synergies The images taken by eyeon may be combined with SAR or other sharper panchromatic images to increase potential segmentation. Some disaster missions, such as flooding, would benefit greatly from data fusion between the multispectral eyeon images and SAR. 3. Proposed Mission Architecture 3.1 Space Segment Technical Concept 3.1.1 Tradeoffs for spacecraft subsystems design In the design process of the subsystems for the spacecraft, few tradeoffs have been considered to opt for the best possible solution, these tradeoffs are, Cost effectiveness Cost is the main drive for the whole mission, and the spacecraft is the most expensive item in it; therefore the cost for the subsystems has been reduced to the bare minimum in order to reduce the Page 12

overall cost of the mission. A balance has been reached between low cost and still adequate quality and reliability. Fulfillment of mission requirements It is possible to sacrifice some of the requirements or to fulfill them partially to gain considerable reduction of complexity and cost of a specific subsystem. In the design the subsystems requirements have been fulfilled in an acceptable and cost effictive way. Physical characteristics (mass, volume) These factors, although not affecting the mission in a direct way, are very crucial in the launcher selection and ongoing operational cost of the mission. They have been considered all the way in the design process. Reliability and availability To increase the reliability one way is to add backups for the specific subsystem. A backup will dramatically increase the reliability of the subsystem; a second backup will even increase it further. Backups add additional cost to the system, increase the overall weight, power consumption and volume. Therefore, a careful design of backup scheme for the various subsystems have been considered. 3.1.2 High-resolution multi-spectral imaging system To fulfil our Mission Requirements and to take all applications into account we addressed in Section (x), we decided on using the JSS-61 space borne scanner from Jena OptronikGmbh. The device is an enhancement of the JSS-56 space borne scanner used for the Rapideye Mission. The main improvements address the resolution and the implementation of an additional band, the panchromatic band. With this additional band (table 3) and the improved resolution we are able to detect and image nearly every common disaster. Another reason for choosing the compact solution, provided by Jena Optronik, is that the device can be used as a standalone solution and no additional time and research has to be spend on developing a sensor. Jena Optronik guarantees the functionality of the whole system for at least 5 years and each subsystem is optimized in a way that the specifications of the sensor can be fulfilled. Nevertheless we have to develop a SSD interface for extending the on-board storage of the sensor about additional 128Gbit storage. This adaption is made for two reasons. First of all we want to extent our maximum track length and secondly we want to provide a redundancy model for the case that the internal storage device fails. With our model, and the interface we want to implement, we are able to either double the track length, our store the data generated by the sensor on the external storage. This is done, because we want to ensure that we can achieve a mission lifetime of at least 5 years. So, in the first stage of our mission we can double the data collection rate and build up our Image Library much faster, and in the second (until end of lifetime) stage we can still provide the basic concept of our mission. We assume that the device delivered by Jena Optronik would fulfill its function for our desired mission time, but we still want to add a redundant system to be sure. In the subsystem we even considered a redundant storage device which is connected to our backup system. So even if our adaption fails, we can provide a redundant system. This is done mainly because we think that our main goal should be to take and send images for the whole lifetime. Page 13

Table 3: Specifications of the camera sensor from Jena Optronik Spectral Bands Pan 450nm 800nm Blue 440nm 510nm Green 520nm 590nm Red 630nm 685nm Red edge 690nm 730nm Near Infrared 760nm 850nm Sensor Parameters Instantaneous Field of View IFOV 0.004297 Field of View FOV 1.73 Total Field of View TFOV 50.02 Potential swath coverage 1431 km Basis length 278.06 km Orbit period 96,62 min Ground track velocity 7.562 km/s Dwell time 0.00596 ms Data rate uncompressed 1.45 Gbit/s @ 12bit depth for all 6 sensors Maximum track length uncompressed / 2:1 DCT 624.91 km / 1249.83 km Type JSS-61 No. of channels 6 No. of pixel per channel 12000 F#/focal length [mm] 248522 GSD @ 600km orbit height 1.5m PAN 4.5m MS Swath width @ 600km orbit height 18km Telescope type Ritchey-Chrétien Sensor type Pushbroom Minimum SNR >50 Digitization depth 12 bit Data compression method DCT Data storage capacity 120 Gbit Data encryption /formatting Triple DES/CCSDS Data downlink interface 4x50Mbps Peak power consumption 180 W (simultaneous image take & downlink) Average power consumption 30 W (including Overall mass 96 kg Overall dimensions imager front-end [mm] 780 x 880 x 1420 Overall dimensions electronic box [mm] 280 x 260 x 420 Page 14

Since the device provided by Jena Optronik is not capable of tilting and off-nadir pointing by itself, a mechanism to tilt the sensor is needed in order to at least achieve an off-nadir pointing of 25. In general there are 3 approaches in use for other missions. The SPOT satellite mission uses a mirror system for pointing the sensor to the desired location. The advantage of this approach is that neither the spacecraft nor the sensor has to be turned. But the disadvantage of the system used for the SPOT satellites is the complexity of the optical mirror and the implementation and adaptation for an already existing sensor like the JSS-61. The effort we assessed it would take us to adapt a mirror system is not related to the investment. A second approach used in the IRS satellite mission provided a Payload Steering Mechanism (PSM) for tilting all on-board sensors up to 27 off-nadir. This approach contains a dedicated plane inside the satellite that can be moved by several motors. Unfortunately this system needs a lot of adaptions for our sensor and also affects the geometry of the satellite. The assumption of a simple geometry design for the eyeon spacecrafts makes it impossible to use this approach. Therefore the already included Attitude Control System is used to tilt the satellite to the desired angle. 3.1.3 Flight Commander The Flight Computer is a high-performance, low-cost single-board spacecraft computer, designed for LEO applications. Two single board computers are used in a redundant configuration. The flight commander is used to control every subsystem and especially take care about the attitude determination and control system. We have chosen a bus topology (CAN bus) for economic reasons, because it needs only one connection to/from the on-board computer. The disadvantage of having only one bus for each device can be compensated by our second redundant flight commander. In case of failure the second flight commander with the redundant bus can stand in for the first bus on the primary flight commander. But nevertheless, each device is redundant either. A flight commander that is capable of several communication protocols like LVDS and CAN was chosen. So we benefit from the low-cost solution but added a redundancy due to the fact, that we double every single important device and connection. Since the flight commander is the most important device, handling all computations for each device, additional shielding against cosmic radiation is added to ensure operation for the desired lifetime and even further. Figure 9 Scheme of topology used for the flight commander 3.1.4 On-board communication For communication purposes a S-band Patch Antenna is used that supports Telemetry and Telecommand data for the Space Mission. The patch antenna produces a hemispherical pattern, with Page 15

good gain along bore sight. The antenna enables us to upload commands to the flight commander which is responsible of the attitude determination and control system. Table 4 S-Band Frequency range used for communication and telecommand purposes Band Frequency Service Direction S 2025-2120 SR, S0, EES Earth Space 2200-2300 SR, S0, EES Space - Earth The primary X-Band transmitter is part of the sensor system, and uses 4x50Mbit/s interfaces to deliver the data to the included X-Band antenna that downlinks the data at 200Mbit/s. To achieve a better performance a second X-Band antenna with a different polarization is used. The secondary X- Band transmitter supports high-speed data return for the additional storage device mounted on the satellite. The flexible architecture allows for switching data rates, modulation, and coding schemes in-orbit while maintaining RF output levels. The data rates provided by the antenna can be adjusted according to our needs. In case the primary X-Band antenna of our sensor fails, we can swap all the data to our secondary antenna and transmit the data with full transmitter capacity of 400Mbit/s to the ground. The time needed to downlink our data is always within the visibility time of the ground station even if the primary antenna fails. 3.1.5 Power subsystem The power system design is one of the very critical parts of the spacecraft design, since the power generation and consumption is affecting the design of all other subsystems in a direct way. The overall power consumption of the spacecraft should be kept as low as possible, provided that it meets the mission power requirements. A typical power system includes a solar array for power generation, a power control module for power regulation and distribution, and a battery system for power storage. In the following section we will discuss these three main components of eyeon power system. 3.1.6 Spacecraft power requirement calculation The design requirements for the power system are that it should provide enough regulated power for every system onboard the spacecraft, and to store enough energy to supply the spacecraft when no power from solar panels is available. Using table 5 the peak power needs for the system has been calculated. 1050.9W is the peak power when all subsystems are drawing their maximum power from the system. In the operation plan it is decided that the peak power will be reduced by not operating the heavily consuming devices at the same time. The plan states that components like the reaction wheels will not be operated all at the same time; instead every maneuver will be implemented in steps to avoid the need for high currents in the system. The backup Sun sensor and the Data Recorder are redundant, they don t need to be considered in the power requirement calculation. Table 5: Power needs of different subsystem components Subsytem max peak [W] rated power [W] JSS-61 180 10 Reaction Wheel 145 16.3 Reaction Wheel 145 16.3 Page 16

Reaction Wheel 145 16.3 Reaction Wheel 145 16.3 Antenna Pointing Mechanism 3.5 3.5 X-Band Transmitter 100 5 S-Band Antenna 10 2 Data Recorder 15 5 Data Recorder 15 5 LEO Flight Commander 10 10 Backup Flight Commander 10 10 GPS Space Receiver 5.5 5.5 Sun Sensor 0.2 0.2 Sun Sensor 0.2 0.2 Star Tracker 20.5 13 Star Tracker 20.5 13 Star Tracker 20.5 13 Interface SSD Extention 10 3 Interface SSD Extention 10 3 Heaters 40 30 Total 1050.9 196.6 The second important item in the power system design is the average power for different subsystems. As explained in table 6, the maximum operating power has been calculated at 359.5W. 15% margin has been added to account for any errors, the final operating power was determined at 420W. Table 6: Operating power of components Subsytem Max. Power [W] Typical Power [W] Minimum Power [W] JSS-61 100 30 5 Reaction Wheel 16.3 16.3 16.3 Antenna Pointing Mechanism 3.5 0 0 X-Band Transmitter 100 5 0 S-Band Antenna 10 5 0 Data Recorder 15 15 5 LEO Flight Commander 10 10 10 Backup Flight Commander 10 10 0 GPS Space Receiver 5.5 5.5 5.5 Sun Sensor 0.2 0.2 0.2 Star Tracker 13 13 13 Star Tracker 13 13 13 Star Tracker 13 13 13 Interface SSD Extention 10 10 0 Heaters 40 30 30 Total 359.5 176 111 Page 17

As can be seen from the tables, the main power requirement is imposed by the primary payload; the multispectral camera. The camera consumes 28% of the maximum operating power. This is still less than the typical value for payload consumption at 40% according to [SMAD]. 3.1.7 Power regulation Figure 8: Power system components An important function of the power system is power regulation. The designed power system is capable of providing regulated voltages with different levels to suit the needs of every subsystem and to exclude additional power conversion modules for individual subsystems as much as possible. The chosen system is Small Satellite Power System from Surrey (figure 8). The regulators provide two levels of voltage, namely 28V and 15V. additional power conditioning is done by different modules separately, employing simple power amplifiers. Another functionality of the designed power system is switching off unneeded systems, to reduce the unnecessary power consumption. 3.1.8 Solar array design Figure 10: Selected Solar cell from emcore The energy source for the spacecraft is the Sun; using solar arrays the power system insures a continuous supply of electrical power to all parts of the spacecraft. The design of solar array depend basically on the power requirement of the system,efficiency of the used array, size of solar array, and the environment the spacecraft is working in (the chosen orbit which determines exposure to sun). To calculate the area of the solar array we need to calculate the power to be generated by it. We start from the operating power requirement of the spacecraft which is found to be 420W as explained in section 3.1.6. Considering the fact that the spacecraft is not always illuminated by the sun (LEO mission), and assuming an 80% efficiency of the solar cells, in addition to the fact that the Page 18

spacecraft will have batteries which need to be charged during the Sun exposure time, the power to be generated from the solar panels has been estimated at 1.8 times the spacecraft operational power. Solar array power = 1.8 x 420W = 750W After calculating the power demand from the solar panel, we compared the available solar cell to choose one of them for eyeon mission. The comparison factors are, high power output (high efficiency), light weight, and acceptable cost. The choices we had were silicon cells and Advanced Triple-Junction (ATJ) InGaP/InGaAs/Ge cells. The silicon cells have low efficiency at 14%, compared to the ATJ cells which have 27% efficiency, but for a much higher cost. The weight for both is almost the same at 1kg/m². ATJ cells from the company emcore have been chosen. Knowing the total demand from the panels as well as the efficiency, we could calculate the required area of the solar array using the formula from [SMAD] Where A a is the area of the array in m² P is the power of the array in Watts e is the cell efficiency 1367 is the solar constant in W/m 2 on earth, approximately the same value at LEO For eyeon mission it is found that This is the total area needed in 100% Sun illumination conditions. To account for the facts that some sides of the spacecraft will be on the shadow, and that the Sun arrays will always hit the solar panels in an angle; the total solar panels area is estimated comparable to the area of RapidEye mission at 10m². The total area of eyeon spacecraft is 10.3m², therefore it was decided that the solar panels will be mounted on the outer surface of the spacecraft, no need for deployable panels. The last design parameter to calculate is the weight of solar array. The solar cells from emcore have a mass density of 85mg / cm², the total array for eyeon spacecraft is 10m² of area. total weight = 10 x 10000 cm²/m² x 85 mg/cm² = 8.5 kg Solar cells degrade considerably with time, a typical value for degradation is 30% after 10 years for GEO missions, and for LEO the value is even larger due to higher ray s rate. The solar panels for eyeon mission are estimated to degrade by 20% in 5 years, but still be able to provide the power required by the spacecraft because of the high design margin added. 3.1.9 Battery system One major limiting factor of spacecraft lifetime is the battery system. The batteries get degraded with time very fast; as a result they lose considerable percentage of their peak performance after few years. In designing battery system for eyeon mission we have considered 3 different battery types, NiH2, NiCd and NiMH. The choice was made favoring the Nickel-Hydrogen (NiH2) batteries listed in Table 7. NiH2 batteries have been specifically designed for space applications, in addition to that they posses Page 19

higher energy density per kilogram compared to NiCd and NiMH counterparts. These batteries have achieved high performance in many space missions in the past, the only drawback of them is the relatively high cost. Table 7: Battery parameters Parameter Value Energy density 80Wh Width 10cm Length 10cm Height 20cm Weight 1.25kg/unit Efficiency 80% Degradation per cycle 20% after 30000 cycles To provide 420 we need six battery units (6 x 80W = 480W). Four additional batteries have been included as a backup. The batteries are arranged in an array to reduce the volume and improve the mass distribution. Battery cells degrade over time. The degradation factor for NiH2 batteries is shown in table 7 at 20% after 30000 cycles. It has to be ensured that throughout the 5 years mission lifetime the batteries will still be able to provide enough energy for the spacecraft. The maximum battery capability for eyeon mission is 10 x 80W = 800W. An eyeon satellite has 15 cycles per day and is set up to fly for 5 years. This results in a total number of cycles of 15 x 365.25 x 5 = 27393.75 for the whole mission. Assuming a full charge/discharge cycle for every orbit and a linear degradation of the batteries (for five years = 27393.75 x 20 / 30000 = 18.2%), the battery capacity at the end of the mission will be (100 18.2 ) x 800 W = 653.9 Wh, which is still well above the mission requirements. Another important limitation brought by the battery system is the temperature range. The battery system is one of the most sensitive components to extreme temperatures changes. Therefore it is very important to consider the temperature range for the chosen batteries. For NiH2 the operational range is 5 C to 20 C. 3.1.10 Power cycle calculation In this section a full cycle of batteries charge/discharge is analyzed, to make sure the power system designed can provide enough power to recharge the batteries to 100% of their capacity, and to ensure that the solar panels will still provide enough energy to operate other spacecraft subsystems. The energy worst case condition for the mission is when the satellite enters the dark region of the orbit and has to download the data stored on board employing the energy supplied from the batteries. Full cycle time = 96.6 minutes Dark period = 37% of full time = 35.5 minutes Page 20

The maximum operating power requirement is 359.5W (table 6), this happens when the spacecraft is downloading the data and taking image at the same time. The energy that is needed to download the data = 360 x 10 /60 = 60Wh. The visibility time at the ground station is 10 minutes; this is the time in which the spacecraft is able to download the data. We assume that the satellite will use all the 10 minutes period. For the rest of the dark time, which is 61 minutes, we need the station keeping rate only, which is 111W as shown in table 6. Station keeping time = dark period time download time Station keeping time = 35.5 10 = 25.5 minutes Station keeping energy in dark = 130W x 25.5 / 60 = 47.18 Wh total energy the system needs to draw from the batteries in the dark side = 60Wh + 47.18Wh = 107.18Wh discharge percentage = 107.18Wh / 800Wh =13% of total capacity For the next cycle the satellite needs to recharge on the sunny side of the orbit. Assuming the smallest side of the spacecraft is totally illuminated by the Sun and using the average value of 275W/m² energy density for the solar cell, the total energy generated by the solar panel can be calculated. energy available during daytime = 61 minutes / 60 x 1.5m x 1.5m x 275W/m² = 620 Wh In case the spacecraft has to take images for 3 minutes, download the data again for 10 minutes and spend the rest of the daytime (61 minutes) in station keeping mode, the total energy needs by the spacecraft is summing to 60Wh + 9.75Wh + 181.13Wh = 250.88Wh. The energy available is 620Wh, the energy available to recharge the batteries is 620Wh 250.88Wh = 367.87Wh. This is more than three times larger than the amount of energy used from the batteries during the dark period of the orbit. This proves that during the charge/discharge cycle the solar panel generates enough power to supply the spacecraft subsystems and recharge the batteries simultaneously. The power system will have considerable losses as heat in wiring; this quantity has been estimated at 3% of the operating power [SMAD], which is covered by the design margin. 3.1.11 Thermal subsystem The Thermal Control System (TCS) is responsible for fulfilment of the temperature requirements of every subsystem and component onboard the spacecraft. The TCS should maintain the components within their operating temperature range in all mission phases and under all light conditions. Designing a thermal system for a LEO spacecraft should take into account the complicated and dynamic nature of LEO orbits. LEO orbits have low altitudes and short orbital periods (96.6 minutes in eyeon case). In LEO orbit a spacecraft sees a small portion of the Earth, and gets exposed to rapid changes in temperature conditions, due to the different time zones and different geographical regions it crosses in short time. The lighting and exposure to sun also varies rapidly in the orbit, nearly half of the orbit period the spacecraft is in daylight, the other half it is in darkness. All these variations affect the temperature of different components of the spacecraft and the TCS has to condition them out. The TCS design is different from other subsystems design in that it is inevitably connected to all other subsystems, affect them and get affected by them. A direct result of this fact is that we have to have a good knowledge about other subsystems before starting designing TCS. Page 21

Figure 11: EyeON heat-environment (Millina F. Diaz-Aguada et al.) 3.1.12 Basic Energy Balance To design a TCS for a spacecraft we need to consider two things, first we need to know the temperature limits of every component, the internal power dissipation, and the worst-case environmental conditions. Second we need to develop a thermal energy balance for the spacecraft. The thermal energy balance determines the internal heat load that the TCS should be able to accommodate. In table 8 we have listed the minimum and maximum temperatures for all components of eyeon spacecraft. The temperature limit for the spacecraft was chosen to be the most tight temperature range among the components. Therefore, the TCS in eyeon mission should maintain the spacecraft within 5 20 C all the time and in all different conditions. Table 8: Temperature limits for different parts Component Lowest temp. [ C] Highest temp. [ C] Data Recorder -20 50 Antenna Pointing -40 60 Sun Sensor -30 65 Star Sensor -40 40 Reaction Wheel -20 50 Power System -20 50 Flight Computer -20 50 GPS Receive -20 50 S-Band Antenna -20 50 Xenon Propulsion System -20 50 Page 22

X-Band Antenna -20 50 Battery 5 20 To develop a thermal energy balance we need to specify all sources of heat affecting eyeon satellite. A spacecraft in LEO orbit is affected by three kinds of heat fluxes, Sun flux, Earth flux and Albedo, these three environmental factors are shown in figure 8. The other source of heat is the internal heat generated by electrical components. The temperature conditions onboard the spacecraft are determined by the sum of the external and internal factors. In the following section we will look at the worst case hot condition and the worst case cold condition, these two cases represent the extreme conditions TCS has to handle. The worst-case hot condition (adopted from Andrew D. Williams) Orbit beta angle is 90. Eclipse duration is zero. The panel with the largest surface area is always nadir facing. The panel with the second largest surface area always faces the sun. Solar flux is 1414 W/m2. Earth IR is 275 W/m2. Albedo coefficient is 0.57. The side reserved for the payload faces space. That side does not radiate heat to space. Figure 9: Worst case hot and cold configuration (Andrew D. Williams) The worst case cold condition is defined by Orbit beta angle is 0. Eclipse duration is 43%. The panel with the smallest surface area is always anti-nadir. Solar flux is 1322 W/m2. The side reserved for the payload is nadir pointing, so there is not an Earth IR or Albedo heat load. Page 23

The TCS function is to make a balance between the heat gained from the environment and the heat dissipated back. Using the energy balance results we can decide whether the spacecraft needs a heater or radiator systems. The equation below is the basic model for the heat energy balance onboard a spacecraft. ε is the emissivity of the spacecraft σ is the Stefan-Boltzmann constant [W/m2 -K4 ] is the radiator surface area [m²] is the average temperature of the spacecraft [K] is the surface area [m²], is the view factor between the spacecraft and the Earth is the intensity of the Earth IR α is the surface solar absorptivity is the area perpendicular to the sun [m²] is the solar heat flux [W/m2] α is the Earth Albedo coefficient is the view factor between the spacecraft and the sunlit Earth is the internal heat generation [W]. For this simple model we consider the temperature of space to be 0 K (actual value is 4 K). In the hot case we assume that all the power consumed onboard the spacecraft has been transferred to heat, in eyeon spacecraft this will be 390W. In the cold case there will be some devices in operation, therefore the lower power is never zero, a typical value used is 50W [Andrew D. Williams]. From the equation above we can find which is the surface area of the radiator required to maintain the satellite within the operating temperature limits in the worst hot case. Using the same equation it is poissble to solve for T S to find the worst cold case temperature. Table 9: Worst hot case and worst cold case parameters Parameter Hot Case Cold Case Eclipse percent 0 0.43 Solar Constant [W/m²] 1414 1322 Albedo Coefficient 0.57 0.18 Earth IR [W/m²] 275 218 Internal Heat [W] 390 50 Temperature Limit [K] 293 278 Area perpen. to Sun [m²] 2.25 1 The value for emissivity used is 0.88, and absorptivity is 0.22, assuming a white paint for the spacecraft. Using the same values as in [Andrew D. Williams], the radiator area was estimated at 1.5 Page 24

m² for the worst hot case to keep the spacecraft below 293 C. A heat-pipes system has been added to the design of the TCS to improve the performance of heat dissipation system. The radiators will be placed on the spacecraft surfaces which are least exposed to Sun or Earth radiations. The heaters selected for eyeon mission are thermostatically guarded polyimide thermo foil working in the range of 22.5 40W. Another item which contributes to the control of the temperature onboard the spacecraft is the coating and insulation of the surface. Surface coating has been added to keep a good balance between the absorbed and dissipated heat. The chosen coating is a multilayer insulator of 10 layers. 3.1.13 Attitude Determination and Control System (ADCS) The Attitude Determination and Control System (ADCS) determines and controls the orientation of the eyeon satellites in space. Attitude control consists of a closed control loop (figure 12) which determines the actual attitude and compares it to the desired attitude. The difference of both is the attitude deviation which is translated into commands for the actuators by the attitude controller. The actuators create torques to change the attitude. This loop is repeated until the desired attitude is achieved. Figure 10: Closed control loop for attitude control (Source: Ley et al. 2009) For attitude determination attitude sensors are used. Each eyeon spacecraft comprises three star trackers of the model ASTRO 10 (figure 13) as well as two fine sun sensors (one for redundancy) (figure 14). These devices are purchased at the company Jena Optronik. Each star sensor consists of an optical head and a separate electronic box to get maximum flexibility for installation of the sensor on the satellite. The star light is projected onto a CCD chip in order to generate a picture of the surrounding star pattern which is compared to an on-board star catalog. The attitude of the spacecraft is then calculated with high accuracy (1.5 arcsec). A sun exclusion angle of 30 and the installation of several sensors in different directions assure a minimization of glare effects from the Earth, Sun and Moon. Peltier elements are used for cooling in order to minimize the radiation damage on the CCD chip. Further specifications of the star tracker ASTRO 10 are given in table 10. Page 25

Table 10: specifications for the ASTRO 10 star tracker from Jena Optronik (source: ASTRO 10 data sheet) Specification Value size optical head (30 baffle) 140 mm Ø x 264 mm size electronic box 150 mm x 145 mm x 75 mm mass head 1180 g for 30 Baffle mass baffle 540 g for 30 Baffle mass electronic box 1360 g focal length 30 mm temperature range -40 C...+40 C field of view 16.7 x 12.5 attitude accuracy 1.5 arcsec (1σ) xy-axes 12 arcsec (1σ) z-axis attitude re-aquisition max. 8 s slew rate up to 1 s-1 power consumption total, peltier cooling off 8 W (at 20 C, at 28 V) total, peltier cooling average 11 W (at 28 V) 15 W (Peltier Cooling max) optical head, peltier cooling off 2 W (at 20 C) optical head, peltier cooling max 5.5 W Figure 13: ASTRO 10 star tracker from Jena Optronik (source: ASTRO 10 data sheet) Page 26

A sun sensor measures the direction to the Sun. The fine sun sensor installed on the eyeon spacecrafts offers high pointing abilities as well as special thermal radiation stability. The light is directed onto the photodiode detector through a slit mask. Sensor specifications are listed in table 11. Table 11: specifications for the Fine Sun Sensor FSS from Jena Optronik (source: Fine Sun Sensor FSS data sheet) Specification Value size 160 mm x 145 mm x 56 mm mass 650 g temperature range -30 C...+65 C field of view 128 accuracy 0.15 (3σ) power consumption 200 mw Figure 11: Fine Sun Sensor FSS from Jena Optronic (source: Fine Sun Sensor FSS data sheet) Attitude control is achieved by using actuators that generate torques which cause changes in the spacecraft s angular momentum. EyeON s attitude control is done by four reaction wheels of the type Smallwheel 200SP (figure 15) offered by Surrey Satellite Technology. A reaction wheel generates an internal torque by accelerating or decelerating. The angular momentum of the change in rotation speed is transferred to the spacecraft causing it to change attitude. The rotation is controlled by an electrical motor in order to achieve different amount of torque. The Smallwheel 200SP achieves a maximum rotation speed of 5000 rpm and can generate a maximum torque of 200 mnm when storing 12 Nms of wheel momentum. This is sufficient to use the reaction wheels as the eyeon camera pointing mechanism. Page 27

Table 12: specifications for the Smallwheel 200SP reaction wheel from Surrey Satellite Technology (Source: Smallwheel 200SP data sheet) Specification wheel momentum wheel torque wheel moment of inertia wheel speed motor torque constant mass Size power consumption temperature range Value max. 12 Nms max. 200 mnm 0.023 kgm2 ± 5000 rpm 0.043 Nm/A 5.2 kg 240 mm diameter, 90 mm height 3.3 W (0 rpm), 16.3 W (5000 rpm), 145 W (max.) -20 C 50 C Figure 12: Reaction Wheel Smallwheel 200SP from Surrey Satellite Technology (Source: Smallwheel 200SP data sheet) External disturbances, such as air drag and solar pressure, always generate torques which act on the spacecraft in the same direction. Therefore they can only be compensated by internal torques for a limited time and additional external actuators need to be used to keep a constant attitude. This is the reason why thrusters are also installed on the eyeon satellites. The Xenon Propulsion System (figure 16) which is sold by Surrey Satellite Technology is designed for in-plane maneuvers for low orbiting satellite missions. It is particularly recommended for high resolution imaging missions by Surrey Satellite Technology. Each eyeon spacecraft will comprise four propulsion systems with one resistojet thruster each in order to cover each direction in space. The included resistojet thruster technology was developed to enhance the specific impulse by further heating the gas before leaving the nozzle with electrical resistance heater elements which results in an increased exhaust velocity. The Xenon Propulsion System will be used for compensating disturbances acting on the spacecraft such as air drag and solar pressure, counteracting the unloading process of the reaction wheels as well as the final deorbiting maneuver. Therefore a capacity of 40 kg of fuel was calculated to achieve the required mission life time. The technical specifications of the system are listed in table13. Page 28

Table 13: specifications of xenon thrusters (source: Xenon thruster data sheet) Specification Value thrust 18 mn resistojet power 2 x 30 W specific impulse 48 s nominal total impulse 5644 Ns warm up 700 s mass 12 kg xenon, 7.26 kg dry mass dimensions 230 x 300 x 300 mm operating temperature -20 C 50 C Figure 13: Xenon hot gas propulsion system (source: Xenon thruster data sheet) 3.1.14 Structure Structure of the spacecraft is designed to meet multifunctional needs of satellite structures addressing the the aggressive structural and acoustic born vibration environments of launch and the stiffness, electrical, thermo-elastic and dimensional stability requirements of in-orbit operation. It is designed in dimensions of 1.5m x 1.5m 1.6m and subsystems are placed into the bus in an optimized way. Camera is placed in the middle of the structure in the center of gravity of satellite bus. Batteries are placed on top since best temparature range is achieved on the top of the bus. Solar arraysare placed on the surface so that the bus does not have wings. Flight computers are placed on the opposite side of flight direction to prevent them to be effected from solar radiation. GPS antennasare placed on top since navigation satellites have higher altitude. Page 29

S-band and X-band antennas are placed under the bus in order to point earth. Star trackers are placed on top, oriented as three dimensional axes. Each pointing direction is 45 degrees away from main axis, which is perpendicular of bus surface. 3 Reaction wheels are also placed as three dimensional axes in a corner. An additional wheel is placed in the middle of these axes 45 degrees away from each of them to be able to compansate in each direction in case one of the main wheels does not function. 3.2 System Budgets 3.2.1 Mass Budget The mass budget of the subsystems is described in table 14. Table 14: Subsystem mass budget Subsytem mass [kg] JSS-61 96 Reaction Wheel 5.2 Reaction Wheel 5.2 Reaction Wheel 5.2 Reaction Wheel 5.2 Antenna Pointing Mechanism 2.7 X-Band Transmitter 4 S-Band Antenna 0.08 Data Recorder 1 Data Recorder 1 LEO Flight Commander 1.5 Backup Flight Commander 1.5 GPS Space Receiver 0.95 Sun Sensor 0.95 Sun Sensor 0.95 Star Tracker 3.08 Star Tracker 3.08 Star Tracker 3.08 Power Conditioning Module 1.65 Battery Charge Regulator 2.25 Battery Charge Regulator 3.25 Interface SSD Extention 0.1 Interface SSD Extention 0.1 NiH2 Battery 1.25 NiH2 Battery 1.25 NiH2 Battery 1.25 NiH2 Battery 1.25 NiH2 Battery 1.25 NiH2 Battery 1.25 Page 30

NiH2 Battery 1.25 NiH2 Battery 1.25 NiH2 Battery 1.25 NiH2 Battery 1.25 Heaters 2 Radiators 1 Structure 180 Solar Array 47.4 Total 390.92 3.2.2 Power Budget Table 15 shows a list of the satellite power budgets. The maximum peak power and the rated power are given. The maximum assumption of the power includes running the image downlink with two satellites without manoeuvre and image taking. The typically consumed power while taking images without manoeuvring is also shown. The last column describes the minimum power consumption when only station keeping is done and the backup system is running. Table 15: Subsystem power budget Subsytem max peak [W] rated power [W] Max. assumption of power [W] typical power consumption [W] minimum power consumptio n [W] JSS-61 180.0 10.0 100 30 5 Reaction Wheel 145.0 16.3 16.3 16.3 16.3 Reaction Wheel 145.0 16.3 16.3 16.3 16.3 Reaction Wheel 145.0 16.3 16.3 16.3 16.3 Reaction Wheel 145.0 16.3 16.3 16.3 16.3 Antenna Pointing 3.5 3.5 3.5 0 0 Mechanism X-Band Transmitter 100.0 5.0 100 5 0 S-Band Antenna 10.0 2.0 10 5 0 Data Recorder 15.0 5.0 15 15 5 Data Recorder 15.0 5.0 0 0 0 LEO Flight 10.0 10.0 10 10 10 Commander Backup Flight 10.0 10.0 10 10 0 Commander GPS Space Receiver 5.5 5.5 5.5 5.5 5.5 Sun Sensor 0.2 0.2 0.2 0.2 0.2 Sun Sensor 0.2 0.2 0.2 0.2 Star Tracker 20.5 13.0 13 13 13 Star Tracker 20.5 13.0 13 13 13 Star Tracker 20.5 13.0 13 13 13 Interface SSD 10.0 3.0 10 10 0 Extention Page 31

Interface SSD 10.0 3.0 0 Extention Heaters 40.0 30.0 40 30 30 Total 1050.9 196.6 368.6 195.1 129.9 3.3 Ground Segment Elements 3.3.1 Ground Stations The duration of the contact and the maximum elevation angle of the satellite depends on how close the ground station is to the satellite's ground track on any given orbit pass [Larson et al. 1999]. Considering 5 satellites, the location is chosen in the vicinity of the area where the subsatellite points of the orbits intersect in order to maximize the operation time of the station.the visibility cone of station can be seen from Figure 14. Figure 14: Visibility Cone of Svalbard Ground Station (STK) The number, locations and variety of users determine the complexity of ground stations [Larson et al., 1999]. In our case, there will be one dedicated ground station. Since our system focuses on providing information to usersall over the world via internet, user positions do not effect the decision criteria of ground stationlocation. 3.3.2 Mission Control Center in Munich A mission control center is decided to be employed in Munich in order to control, monitor the health, track the position of space segment andto determine spacecraft attitude. Main responsibilities are stated below. Spacecraft monitoring Anomaly handling Spacecraft operations Data acquisition and distribution Mission planning Page 32

Navigation and Flight Dynamics[Ley et al. 2009] 3.3.3 Ground Station in Svalbard/Norway There will be a new ground station in Svalbard (78 N 16 E) for data acquisition and transfer to mission control center. Type of ground stations chosen to be dedicated since our constellation of satellites can be operated sequentially and it is cost effective in terms of operation costs.exact location is not specified due to the fact that site selection requires more detailed analysis. It is proven from calculations in the following part that it is possible to download all data for one period from our operation center stationin visibility time. Additionally, visibility times are analyzed via STK software. As can be seen from Figure 18, our ground station have full access to all satellites and operates full time in a day. Figure 15: Visibility times of the eyeon satellites at the Svalbard ground station 3.3.4 Ground Segment Architecture The eyeon system is designed in a way, that it can fulfil costumer s needs within 2 days for registered users and within 4 days for unregistered users. The difference for both users groups is beside the permitted delivery date, the systems provided to order data. For the unregistered user it s only possible to order imagery data out of the library that has already been built upon that time, while the registered user is provided with an emergency hotline and the ability to order data accordingly to his needs. Therefore, eyeon is going to adapt its orbits and viewing angles according to the specified region between -70 and +70 latitude. All orders are contributing to the tasking of further constellation changes and managed in an Emergency Ticketing System and Operational Planning System operated at Munich. The tasking is executed by the systems of the Ground Segment described in Section 1.5. Page 33

Figure 16: High level ground segment architecture 3.3.5 Up/downlink Required uplink and downlink frequency bands are defined in table 16. The corresponding communication bands are chosen according to the frequency band designation as S-band for telecommand uplink and telemetry downlink; and X-band for image downlink. Respective bit rates are also summarized in table 16. Table 16: Up and downlink frequency domain and bit rates (Ley et al. 2009) Data Description Frequency Domain Useful Bit Rate TT&C uplink S-band 2025 2120 MHz 2M bps TT&C downlink S-band 2200 2300 MHz 4096 bps Global Data Stream downlink X-band 8025-8400 MHz 320 Mbps In order to determine whether the visibility time is enough to download images, the following formula and visibility time report from STK software are used. Since the total on board data storage capacity is 248 Gb for one period and total downlink rate is 600 Mb/s for 2antennastherequired time for downloading images is computed as 6.8 min. Considering that the camera will be used only for illuminated conditions,the data amount that should be stored will be smaller, so the actual transfer time will be less. From STK report, it can be seen that minimum visibility time of 8 min is more than sufficient for the download process. Assuming that telecommand and telemetry data are much smaller than image storage, it is feasible to upload and download from each satellite via our ground station. i e ne essa t n a n a ata st a e ata n in ate Page 34

3.3.6 Data Handling Data delivery is based on demand from a near-real-time processing system. The user will request data and be delivered via file transfer to allow remote access of data. Thus network functions and protocols which are compatible with TCP/IP are required. If required it is possible to distribute data via CD/DVD or similar storage options. The system also allows users to transfer command requests as input to our mission planning system. After receiving all requests, mission planning software generates a final schedule. The user has the possibility to access raw telemetry and auxiliary data, which will be kept in a long term archive. There will also be a configuration control function to enable users to retrieve data. 3.4 Mission Phases Summary Mission planning Mission Planning phase involves research, design and verification of all mission elements. The phase ends after the Launch Readiness Review. Definitions of reviews will follow upon proposal acceptance. Launch Launch Phase includes spacecraft launch to orbital insertion. All control is centered at the launch operation center. Following a successful orbital insertion, the phase will transfer to Operations Phase. Operation phase Operation Phase follows commissioning of the satellite. A period of 4-5 days will follow the orbital insertion, during which the camera and sensors are calculated. The ground station will also verify its uplink/downlink capability at this time. The operations phase ends at decommissioning prior to deorbit burn. End of Mission (EOM), Deorbiting At the end of mission, the required 100 m/s delta-v will be expended to send the satellites into a max. 25 year deorbit trajectory. This action complies with the ESA Doc. No: EOL-OHB-ES-001, which conaints strategies for de-orbiting for Debris mitigation. 3.5 Launcher The selected launch system is the Rockot Launch System, delivered and Operated by Eurockot Launch Services, a cooperation between ESA and Khrunichev State Research and Production Space Center. The required Delta-V to reach the desired orbit was estimated to be 7.55 km/s. Several launch configurations were considered, including single launch, pair launch and all satellites in one launch system. The single launch configuration was eventually selected in order to reduce complexity in the satellite platform and because the satellite mass of 430 kg was too large to launch two simultaneously. Concept Risk Complexity ΔV Cost Individual Launch Pair Launch Low Low (+1) Low (launch and position 1) Medium (launch and position 2 simultaneously) Low Medium/Low All in one shot High High (launch and position 5 simultaneously) High Low High Medium Page 35

The selected launch center is Baikonur Cosmodrome. The 45 degree latitude at Baikonur reduces the Delta-V necessary to reach the 97.7 degree inclination for each eyeon satellite. The launch azimuth was calculated to be -11.04, very close to the allowed launch azimuth of -10.9 degrees. The discrepancy in launch azimuth translates larger Delta-V required for launch (Soyuz Handbook). This fact was considered in the Delta-V budget for launch. Phase Launch Correction(2%) Pointing Station Keeping Deorbit (EOL) Total: Total With Launch: Constellation Total: ΔV 7559 m/s 150 m/s 30m/s/year 6 m/s/year 100 m/s 315 m/s 7989 m/s 39195 m/s The estimated Delta-V provided by the launch system was 8.3 km/s to 700km 97.7 degree circular orbit from Baikonur. This value demonstrates feasibility of the Rockot Launch System for use in the eyeon mission. Each eyeon satellite will be launch sequentially with two week separation periods. After 10 weeks, the eyeon constellation will be flying. Other Launch sites will be considered as necessary to complete the launch schedule. 4. Programmatic Elements 4.1 Product Description and Delivery Our system will provide Level 1 to Level 3 data to usersin less than 4 daysand even less for emergency situations. Data will be searched via our user interface enabling users to search our database for interested applications. EyeON imagery products offer a variety of options for accurate and timelyimagery. Imagery products at three processing levels will be available. (1) Basic Imagery with the least amount of processing (geometrically raw), designed for customers desiring to process imagery into a useable form themselves (e.g. research facilities). (2) Standard imagery with radiometric and geometric corrections, delivered in a map projection. (3) Orthorectified imagery with radiometric, geometric, and topographic correction, anddelivered in a map projection. (4) Orthorectified imagery with all correction (like 3) but additionally added information after supervision of one of eyeon s imagery specialists. In addition to processing levels, eyeon offers imagery products in five product options. Page 36

(1) Black & white (panchromatic) products enable superior visual analysis. (2) Multispectral products cover the visible and NIR wavelengths and are ideal formultispectral analysis. (3) Bundle (black & white and multispectral) products. (4) Color (3-band natural color or color infrared) products that combine the visual information of three visible bands with the spatial information of the panchromatic band. (5) Pan-sharpened (4-band) that combines the visual information of all four multispectral bands with the spatial information of the panchromatic band. EyeON provides customers with flexibility in product levels, ordering options, delivery options, and licensing options. EyeON Satellite Constellation: All eyeon satellites collect high resolution imagery with the same multispectral sensors and resolutions. Product Ordering: The eyeon image library allows customers to order previously collected imagery. If the customer s area of interest is not covered in the image library, eyeon will task the image collection in one of two commercial tasking options. Product Delivery: EyeON delivers its imagery products to customers in industry by using standard image formats and media. Customers may receive their products via storage devices shipped with commercial delivery services or electronically via ftp (pull). Product Levels: EyeON offers imagery products in several processing and accuracy levels to suit customers needs. Product Options: EyeON offers products in a variety of multispectral options. Product Naming: Product file naming conventions provide information about the acquisitionof the eyeon imagery. Product Licensing: EyeON licenses its imagery in a flexible manner to ensure that all personnel who need access to the imagery may use it. Image Support Data: The metadata provided with our imagery products provides all the information needed to analyze and process imagery to the customer s specifications. There are two commercial tasking options for our imagery products: Standard and Priority. The tasks available for a single costumer depend on the predefined status of the costumer. When preparing its collection plan, eyeon considers several factors, including order priority, date an order was received, the customer-specified collection window, and the cloud cover forecast. In rare instances, eyeon may pre-empt some orders due to collection efficiency and/or satellite calibration and maintenance. For Standard and Priority Tasking, the customer will be contacted prior to the end of the collection window to cancel the order or extend the collection window if the original tasking order was unsuccessful. Page 37

Physical feasibility assesses the number of times that the eyeon constellation has physical access to the target based upon the parameters the customer provides. Items that affect physical feasibility include off-nadir angle (wider angles will have more accesses than narrow angles), latitude (the eyeon constellation has increased access to locations at higher latitudes), collection windows (the larger the collection window, the more access the satellites will have), and cloud cover forecast. Note that orders over a 25 degree off-nadir angle will require a special review. Competitive feasibility assesses eyeon s ability to collect your order based upon other orders in the system. Items that affect competitive feasibility include orders already in the system and orders that have a higher relative priority. Standard Tasking Order placed at least 48 hours before start collect date. Customer specifies length of collection window up to 365 days. Collection window is subject to feasibility. DigitalGlobe suggests a 90 day collection window to ensure enough time to collect imagery that meets your specifications. Larger areas will require a longer collectionwindow. Unlimited number of collection attempts within the customer specified collection window, depending on off-nadir angle, latitude, competition, and length of collection window. A longer collection window will result in more physical collection opportunities and increased likelihood of a successful collect. Standard image processing, where the defined Standard represents 3-band visible pansharpend imagery. Priority Tasking Order placed at least 24 hours before the start collect date. Collection window of 1-14 days. A single collection attempt. Basic and Standard Imagery only. Rush image processing. Post-processed and thematically adapted Image Library In addition to tasking the satellite, customers may order imagery products directly out of the eyeon image library. Customers may define their order polygons in several ways (table 17). Image library orders receive the Standard level image processing, unless ordered as a priority image library order. Table 17: Description of products Product Level Processing Absolute Accuracy CE90% RMSE Page 38 Geographic Availability Basis Imagery Sensor Corrected (RAW) 23m 14m -70 to 70 latitude Standard Imagery Georectified 23m 14m -70 to 70 latitude Ortho 1:25000 Orthorectified 25.4m 15.4m -70 to 70 latitude Post-processed 1:25000 Thematically adapted 25.4m 15.4m -70 to 70 latitude

Resolution for multispectral imagery ranges from 4.50 m (at nadir) to 5.12m (25 off-nadir looking). The image is resampled to a coordinate system defined by the ideal basic imagery camera model. The resulting GSD varies over the entire product as a function of the attitude & ephemeris during the imaging process. Basic imagery products are not available with pan-sharpening. The radiometric corrections applied to this product include: Relative radiometric response between detectors removes differences in imagery due to sensitivity variations between pixels Non-responsive detector fill fills in null values on imagery due to detectors that are no longer collecting data Conversion for absolute radiometry calibrates overall detector response from known radiometric signals The sensor corrections account for: Internal detector geometry combines the six digital chip assemblies into a virtual array Optical distortion corrects image distortion caused by sensor optics Scan distortion corrects for distortions caused by slew and scan rate Line-rate variations corrects for variations in the panchromatic scan rate Registration of the multispectral bands all multispectral bands the panchromatic and multispectral bands are not registered Physical Structure: Basic Imagery products are delivered as scenes. In length, a scene will be approximately 18 km. In width, a scene will be the full strip width (12,000 pan pixels). The area that this width represents on the ground depends on the collection parameters of the scene (off-nadir angle, orientation of collection, etc). A scene has an approximate area of 324 km 2 (18 km by 18 km) at nadir. Figure 17: visualisation of the size of a scene Page 39

4.2 ROM Cost Estimate 4.2.1 Ground Segment Total cost of Ground Segment establishment and operations are estimated as 3 759 00 for 5 years. Details of estimation are stated in the Table 18 (Angepat, 2005). Table 18: Cost Estimation of Ground Segment Item Cost Antenna Systems 1 674 000 Shipping and Installation 89 000 Additional Options 93 000 Essential on-site spares 65 000 Baseband Equipment 154 000 Radar Dome 122 000 Site Acquisition Preparation 116 000 Power Grid Layout 120 000 Remote Access Site and Equipment (5 years) 257 000 Labor and Maintenance from Remote Access Site (5 years) 1 064 000 Frequency Fee (5 Years) 5 000 Total 3 759 000 4.2.2 Subsystems Table 19 lists all estimated costs for the subsystem components. Table 19: Cost estimation for subsystems Subsytem costs [$] JSS-61 10000000 Reaction Wheel 136850 Reaction Wheel 136850 Reaction Wheel 136850 Reaction Wheel 136850 Antenna Pointing Mechanism 443900 X-Band Transmitter 443900 S-Band Antenna 200000 Data Recorder 266800 Data Recorder 266800 LEO Flight Commander 1000000 Backup Flight Commander 1000000 GPS Space Receiver 182900 Sun Sensor 50000 Sun Sensor 50000 Star Tracker 500000 Star Tracker 500000 Star Tracker 500000 Power Conditioning Module 50000 Page 40

Battery Charge Regulator 50000 Battery Charge Regulator 50000 Interface SSD Extention 100000 Interface SSD Extention 100000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 NiH2 Battery 20000 Heaters 50000 Radiators 30000 Structure 6000000 Solar Array 1000000 Total 23581700 Page 41

References 1) ASTRO 10 data sheet, http://www.jena-optronik.de/en/aocs/astro10.html (accessed on 26. June 2012) 2) Fine Sun Sensor FSS data sheet, http://www.jena-optronik.de/en/aocs/fss.html (accessed on 26. June 2012) 3) FAO(Food and Agriculture Organization) of the United Nations 2005, Mapping global, urban and rural population distributions, by M. Salvatore, F. Pozzi, E. Ataman, B. Huddleston and M. Bloise, Environmental and Natural Resources Working Paper No. 24, Rome 4) FAO & IIASA, 2006. Mapping biophysical factors that influence agricultural production and rural vulnerability, by H. van Velthuizenet. al., Environment and Natural Resources Series No. 11. Rome. 5) Gene Dial, Howard Bowen, Frank Gerlach, JacekGrodecki, Rick Oleszczuk, 2003, IKONOS satellite, imagery, and products, Remote Sensing of Environment 88, 23 36, doi:10.1016/j.rse.2003.08.014 6) Gilmore, D.G., Spacecraft Thermal Control Handbook, Volume 1: Fundamental Technologies, Second Edition, Reston, VA: AIAA, 2002 7) Millan F. Diaz-Aguada et al, Small Satellite Thermal Design, Test and Analysis, The University of Texas at Austin 8) http://www.eoportal.org/directory/pres_ikonos2.html 9) http://www.eoportal.org/directory/pres_rapideyesatelliteconstellation.html 10) http://www.ferris.edu/faculty/burtchr/sure382/lessons/lesson_3.pdf 11) http://www.isprs.org/proceedings/xxxiii/congress/part7/1609_xxxiii-part7.pdf (accessed on 26. June 2012) 12) Kaluza P., Kölzsch A., Gastner M. T., and Blasius B., The complex network of global cargo ship movements, 1001.2172 (January 13, 2010),http://arxiv.org/abs/1001.2172 (accessed on 25. June 2012) 13) Ley W., Wittmann K., Hallmann W., Handbook of Space Technology, John Wiley & Sons, Ltd. 2009, ISBN: 978-0-470-69739-9 14) Larson W. L. and Wertz J. R., Space Mission Analysis and Design, 1999. 15) Münchener Rückversicherungs-Gesellschaft, Geo Risk Research 2012,NatCatSERVICE, http://www.munichre.com/app_pages/www/@res/pdf/natcatservice/annual_statistics/201 1/2011_mrnatcatservice_natural_disasters2011_worldmap_en.pdf?2 (accessed on 25. June2012) 16) Nirupama, Slobodan P. Simonovic (2002), Role of Remote Sensing in Disaster Management, ICLR Research, Paper Series No. 21 Page 42

17) QUICKBIRD Data Sheet, http://www.digitalglobe.com/downloads/quickbird-ds-qb-web.pdf (accessed on 25. June 2012) 18) Topan H., Büyüksalih G., Jacobsen K., Information Contents of High Resolution Satellite Images 19) vanwesten C., Remote Sensing for Natural Disaster Management, International Archives of Photogrammetry and Remote Sensing. Vol. XXXIII, Part B7. Amsterdam 2000 20) Williams A. D., Issues and Implications of the Thermal Control System on Responsive Space Missions, 20th Annual AIAA/USU Conference on Small Satellites 21) www.minco.com/heater_config/ 22) Xenon Propulsion System data sheet, http://www.sst-us.com/shop/satellitesubsystems/propulsion-systems/xenon-propulsion-system (accessed on 28.06.2012) 23) Angepat S, Ground Station Selection and Feasibility Analysis, Carleton University Spacecraft Design Project, 2005 24) Wood, Lloyd First Annual CCSR Research Symposium (CRS 2011), Centre for Communication Systems Research, 30 June 2011 25) ESA End-Of-Life De-Orbit Strategies Document Number: EOL-OHB-ES-001 03.07.2002 26) R. Janovsky, END-OF-LIFE DE-ORBITING STRATEGIES FOR SATELLITES, DLR Luft- und Raumfahrtkongress 2002 27) Soyuz Users Manual, ST-GTD-SUM-01 - ISSUE 3 - REVISION 0 - APRIL 2001 28) AGI STK 9 Pro Tutorial Page 43

Annexes Satellite Visibility Times from Svalbard Ground Station STK Report 27 Jun 2012 22:38:57 Educational Use Only Facility-Facility7-To-Satellite-Satellite6711, Satellite-Satellite6721, Satellite-Satellite6731, Satellite- Satellite6741, Satellite-Satellite6751: Access Summary Report Facility7-To-Satellite6711 -------------------------- Access Start Time (UTCG) Stop Time (UTCG) Duration (sec) ------ ------------------------ ------------------------ -------------- 1 25 Jun 2012 10:21:22.172 25 Jun 2012 10:33:34.797 732.625 2 25 Jun 2012 11:57:41.156 25 Jun 2012 12:10:32.480 771.324 3 25 Jun 2012 13:33:45.401 25 Jun 2012 13:46:43.701 778.300 4 25 Jun 2012 15:09:33.927 25 Jun 2012 15:22:23.746 769.820 5 25 Jun 2012 16:45:08.090 25 Jun 2012 16:57:51.926 763.837 6 25 Jun 2012 18:20:36.116 25 Jun 2012 18:33:25.027 768.911 7 25 Jun 2012 19:56:14.202 25 Jun 2012 20:09:12.063 777.861 8 25 Jun 2012 21:32:21.732 25 Jun 2012 21:45:14.864 773.132 9 25 Jun 2012 23:09:14.585 25 Jun 2012 23:21:32.483 737.898 10 26 Jun 2012 00:47:00.271 26 Jun 2012 00:58:05.220 664.949 11 26 Jun 2012 02:25:32.027 26 Jun 2012 02:34:58.262 566.235 12 26 Jun 2012 04:04:19.848 26 Jun 2012 04:12:26.004 486.156 13 26 Jun 2012 05:42:36.383 26 Jun 2012 05:50:44.062 487.679 14 26 Jun 2012 07:20:02.743 26 Jun 2012 07:29:32.093 569.350 15 26 Jun 2012 08:56:54.994 26 Jun 2012 09:08:02.731 667.737 Page 44

Global Statistics ----------------- Min Duration 12 26 Jun 2012 04:04:19.848 26 Jun 2012 04:12:26.004 486.156 Max Duration 3 25 Jun 2012 13:33:45.401 25 Jun 2012 13:46:43.701 778.300 Mean Duration 687.721 Total Duration 10315.813 Facility7-To-Satellite6721 -------------------------- Access Start Time (UTCG) Stop Time (UTCG) Duration (sec) ------ ------------------------ ------------------------ -------------- 1 25 Jun 2012 10:01:13.655 25 Jun 2012 10:09:07.870 474.215 2 25 Jun 2012 11:38:53.879 25 Jun 2012 11:47:52.052 538.173 3 25 Jun 2012 13:15:54.046 25 Jun 2012 13:26:32.562 638.515 4 25 Jun 2012 14:52:31.137 25 Jun 2012 15:04:32.333 721.196 5 25 Jun 2012 16:28:52.436 25 Jun 2012 16:41:39.284 766.849 6 25 Jun 2012 18:04:59.144 25 Jun 2012 18:17:57.833 778.689 7 25 Jun 2012 19:40:50.313 25 Jun 2012 19:53:42.062 771.748 8 25 Jun 2012 21:16:26.572 25 Jun 2012 21:29:10.984 764.412 9 25 Jun 2012 22:51:54.709 25 Jun 2012 23:04:42.302 767.593 10 26 Jun 2012 00:27:29.714 26 Jun 2012 00:40:26.521 776.807 11 26 Jun 2012 02:03:30.903 26 Jun 2012 02:16:26.464 775.561 12 26 Jun 2012 03:40:15.158 26 Jun 2012 03:52:41.348 746.190 13 26 Jun 2012 05:17:51.813 26 Jun 2012 05:29:11.008 679.195 14 26 Jun 2012 06:56:17.220 26 Jun 2012 07:05:59.453 582.232 15 26 Jun 2012 08:35:05.906 26 Jun 2012 08:43:19.355 493.449 Page 45

Global Statistics ----------------- Min Duration 1 25 Jun 2012 10:01:13.655 25 Jun 2012 10:09:07.870 474.215 Max Duration 6 25 Jun 2012 18:04:59.144 25 Jun 2012 18:17:57.833 778.689 Mean Duration 684.988 Total Duration 10274.824 Facility7-To-Satellite6731 -------------------------- Access Start Time (UTCG) Stop Time (UTCG) Duration (sec) ------ ------------------------ ------------------------ -------------- 1 25 Jun 2012 11:14:40.435 25 Jun 2012 11:24:51.349 610.913 2 25 Jun 2012 12:53:26.947 25 Jun 2012 13:01:59.336 512.389 3 25 Jun 2012 14:32:04.214 25 Jun 2012 14:39:52.527 468.313 4 25 Jun 2012 16:09:53.124 25 Jun 2012 16:18:32.451 519.327 5 25 Jun 2012 17:46:58.356 25 Jun 2012 17:57:17.522 619.166 6 25 Jun 2012 19:23:38.301 25 Jun 2012 19:35:26.318 708.016 7 25 Jun 2012 21:00:01.838 25 Jun 2012 21:12:42.831 760.993 8 25 Jun 2012 22:36:10.893 25 Jun 2012 22:49:09.259 778.367 9 26 Jun 2012 00:12:04.659 26 Jun 2012 00:24:58.345 773.686 10 26 Jun 2012 01:47:43.172 26 Jun 2012 02:00:28.620 765.449 11 26 Jun 2012 03:23:11.867 26 Jun 2012 03:35:58.489 766.623 12 26 Jun 2012 04:58:44.395 26 Jun 2012 05:11:39.937 775.543 13 26 Jun 2012 06:34:39.767 26 Jun 2012 06:47:36.929 777.162 14 26 Jun 2012 08:11:15.714 26 Jun 2012 08:23:48.989 753.275 15 26 Jun 2012 09:48:43.245 26 Jun 2012 10:00:00.000 676.755 Page 46

Global Statistics ----------------- Min Duration 3 25 Jun 2012 14:32:04.214 25 Jun 2012 14:39:52.527 468.313 Max Duration 8 25 Jun 2012 22:36:10.893 25 Jun 2012 22:49:09.259 778.367 Mean Duration 684.398 Total Duration 10265.974 Facility7-To-Satellite6741 -------------------------- Access Start Time (UTCG) Stop Time (UTCG) Duration (sec) ------ ------------------------ ------------------------ -------------- 1 25 Jun 2012 10:53:46.992 25 Jun 2012 11:06:45.582 778.590 2 25 Jun 2012 12:30:09.184 25 Jun 2012 12:42:53.268 764.084 3 25 Jun 2012 14:07:20.621 25 Jun 2012 14:19:15.442 714.822 4 25 Jun 2012 15:45:24.744 25 Jun 2012 15:55:53.642 628.898 5 25 Jun 2012 17:24:07.720 25 Jun 2012 17:32:55.843 528.123 6 25 Jun 2012 19:02:50.319 25 Jun 2012 19:10:39.849 469.531 7 25 Jun 2012 20:40:48.583 25 Jun 2012 20:49:13.614 505.031 8 25 Jun 2012 22:18:00.032 25 Jun 2012 22:28:01.392 601.360 9 25 Jun 2012 23:54:43.540 26 Jun 2012 00:06:18.369 694.829 10 26 Jun 2012 01:31:09.722 26 Jun 2012 01:43:44.273 754.552 11 26 Jun 2012 03:07:21.375 26 Jun 2012 03:20:18.794 777.419 12 26 Jun 2012 04:43:17.918 26 Jun 2012 04:56:13.196 775.278 13 26 Jun 2012 06:18:58.948 26 Jun 2012 06:31:45.338 766.390 14 26 Jun 2012 07:54:28.615 26 Jun 2012 08:07:14.168 765.553 15 26 Jun 2012 09:29:59.156 26 Jun 2012 09:42:53.104 773.948 Page 47

Global Statistics ----------------- Min Duration 6 25 Jun 2012 19:02:50.319 25 Jun 2012 19:10:39.849 469.531 Max Duration 1 25 Jun 2012 10:53:46.992 25 Jun 2012 11:06:45.582 778.590 Mean Duration 686.560 Total Duration 10298.406 Facility7-To-Satellite6751 -------------------------- Access Start Time (UTCG) Stop Time (UTCG) Duration (sec) ------ ------------------------ ------------------------ -------------- 1 25 Jun 2012 10:38:16.724 25 Jun 2012 10:51:05.081 768.357 2 25 Jun 2012 12:13:48.957 25 Jun 2012 12:26:33.105 764.148 3 25 Jun 2012 13:49:17.419 25 Jun 2012 14:02:08.186 770.767 4 25 Jun 2012 15:24:59.335 25 Jun 2012 15:37:57.887 778.552 5 25 Jun 2012 17:01:13.902 25 Jun 2012 17:14:03.213 769.311 6 25 Jun 2012 18:38:15.859 25 Jun 2012 18:50:23.181 727.322 7 25 Jun 2012 20:16:10.777 25 Jun 2012 20:26:58.666 647.889 8 25 Jun 2012 21:54:48.531 25 Jun 2012 22:03:56.146 547.615 9 25 Jun 2012 23:33:34.476 25 Jun 2012 23:41:31.781 477.305 10 26 Jun 2012 01:11:42.137 26 Jun 2012 01:19:58.037 495.900 11 26 Jun 2012 02:49:00.862 26 Jun 2012 02:58:46.577 585.716 12 26 Jun 2012 04:25:48.668 26 Jun 2012 04:37:10.675 682.007 13 26 Jun 2012 06:02:17.904 26 Jun 2012 06:14:45.648 747.744 14 26 Jun 2012 07:38:32.394 26 Jun 2012 07:51:28.370 775.976 15 26 Jun 2012 09:14:31.890 26 Jun 2012 09:27:28.458 776.568 Page 48

Global Statistics ----------------- Min Duration 9 25 Jun 2012 23:33:34.476 25 Jun 2012 23:41:31.781 477.305 Max Duration 4 25 Jun 2012 15:24:59.335 25 Jun 2012 15:37:57.887 778.552 Mean Duration 687.678 Total Duration 10315. Page 49

Detailed Satellite Bus Architecture 10x 1.25 kg ~$200000 2 kg ~$50000 NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh 5.2 kg ~$136.850 Small Reaction Wheel Wheel Momentum 12 Nms Wheel Torque 200 mnm 2.25 kg ~$50.000 Power generated by solar panel (front/right) 28V power 145W peak 16.3W mean Power generated by solar panel (rear/left) 2.25 kg ~$50.000 28V power 180W peak 10W mean Heater NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh NiH² Battery 28V 80Wh 6x BCR Battery Charge Regulator 6x 80W 10x 28V power switches 6x BCR Battery Charge Regulator 6x 80W 10x 28V power switches JSS-61 Spaceborne MS Sensor Data Recorder 120 Gbit Up to 200 Mbit/s (4x50 Mbit/s) Earth Facing Antenna 3,08 kg ~$500.000 3,08 kg ~$500.000 3,08 kg ~$500.000 Star Tracker Star Tracker Star Tracker 950g ~$100,000 GPS Antenna GPS Antenna E-Box E-Box E-Box Sun Sensor Sun Sensor (Backup) 28V power 145W peak 16.3W mean 12V power 200mW peak 200mW mean 28V power 5.5W peak 5.5W mean Small Reaction Wheel Wheel Momentum 12 Nms Wheel Torque 200 mnm 12V power 200mW peak 200mW mean 28V power 20.5W peak 13W mean 28V power 20.5W peak 13W mean GPS Space Receiver 950g ~$182,900 5.2 kg ~$136.850 CAN TM/TC 28V power 20.5W peak 13W mean CAN TM/TC optical RS485 CAN TM/TC optical optical 28V power 145W peak 16.3W mean RS485 28V power 15W peak 5W mean Small Reaction Wheel Wheel Momentum 12 Nms Wheel Torque 200 mnm 2*5V power 10W peak 10W mean Interface SSD Extention Data LVDS Backup Flight Commander 8x LVDS input 8x LVDS output 8x RS485/MLVDS 4x optical input 4x optical output 1.5 kg ~$1 Mio. 5.2 kg ~$136.850 CAN TM/TC < 50g ~$100.000 Data LVDS PCM Power Conditioning Module 28 x 28V power switches 16 x 5V power switches 1.65 kg ~$50.000 Data LVDS CAN TM/TC 1.5 kg ~$1 Mio. 2*5V power 10W peak 10W mean Data LVDS LEO Flight Commander 8x LVDS input 8x LVDS output 8x RS485/MLVDS 4x optical input 4x optical output CAN TM/TC CAN TM/TC Data LVDS Data LVDS 15 to 50V power 100W peak 55W mean Interface SSD Extention 28V power 15W peak 5W mean Data Recorder 128 Gbit 5Gbit/s input 1 kg $266800 Data Recorder 128 Gbit 5Gbit/s input < 50g ~$100.000 Data CAN TM/TC 1 kg $266800 Data LVDS Data LVDS CAN TM/TC 28V power 10W peak 2W mean Data LVDS S-Band Up/ Down < 80g $50.000 CAN TM/TC X-Band Down Up to 400 Mbit/s 4 kg $443,900 96 kg ~$10 Mio. Space Facing Antenna Earth Facing Antenna 28V power 3.5W peak Antenna Pointing Mechanism Up to 400 Mbit/s CAN TM/TC 2.7 kg $443,900 28V power 145W peak 16.3W mean Small Reaction Wheel Wheel Momentum 12 Nms Wheel Torque 200 mnm Earth Facing Antenna 5.2 kg ~$136.850 Data wire Power wire Optical wire Infobox Scale 1:10