imaging satellite mission: eyeon

Transcription

1 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

2 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

3 Table of Contents Executive Summary Objectives and Requirements Objectives Mission Requirements FunctionalRequirements Operational Requirements Constraints Background Existing Imaging Missions Characterization of Area of Interest Applications Retrieval Algorithms Mission Assumption and Technical Requirements Mission Characteristics Mission Simulations and Technical Requirements Complementary Data and Synergies Proposed Mission Architecture Space Segment Technical Concept Tradeoffs for spacecraft subsystems design High-resolution multi-spectral imaging system Flight Commander On-board communication Power subsystem Spacecraft power requirement calculation Power regulation Solar array design Battery system Power cycle calculation Thermal subsystem Basic Energy Balance Attitude Determination and Control System (ADCS) Structure Page 2

4 3.2 System Budgets Mass Budget Power Budget Ground Segment Elements Ground Stations Mission Control Center in Munich Ground Station in Svalbard/Norway Ground Segment Architecture Up/downlink Data Handling Mission Phases Summary Launcher Programmatic Elements Product Description and Delivery ROM Cost Estimate Ground Segment Subsystems References Annexes Page 3

5 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 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 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 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 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

6 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 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) m (25 off-nadir) panchromatic 2.44m (nadir) m (25 off-nadir) multispectral 1 m pan-sharpened spectral resolution PAN, VIS, NIR VIS, Red Edge, NIR PAN, VIS, NIR Page 5

7 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 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

8 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

9 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

10 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

11 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

12 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 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

13 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 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

14 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 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

15 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 Field of View FOV 1.73 Total Field of View TFOV Potential swath coverage 1431 km Basis length km Orbit period 96,62 min Ground track velocity km/s Dwell time ms Data rate uncompressed bit depth for all 6 sensors Maximum track length uncompressed / 2:1 DCT km / km Type JSS-61 No. of channels 6 No. of pixel per channel F#/focal length [mm] km orbit height 1.5m PAN 4.5m MS Swath 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

16 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 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 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

17 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 SR, S0, EES Earth Space 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 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 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 W 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 Reaction Wheel Reaction Wheel Page 16

18 Reaction Wheel Reaction Wheel Antenna Pointing Mechanism X-Band Transmitter S-Band Antenna 10 2 Data Recorder 15 5 Data Recorder 15 5 LEO Flight Commander Backup Flight Commander GPS Space Receiver Sun Sensor Sun Sensor Star Tracker Star Tracker Star Tracker Interface SSD Extention 10 3 Interface SSD Extention 10 3 Heaters Total 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 Reaction Wheel Antenna Pointing Mechanism X-Band Transmitter S-Band Antenna Data Recorder LEO Flight Commander Backup Flight Commander GPS Space Receiver Sun Sensor Star Tracker Star Tracker Star Tracker Interface SSD Extention Heaters Total Page 17

19 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] 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 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 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

20 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 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 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

21 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 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 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 x 5 = for the whole mission. Assuming a full charge/discharge cycle for every orbit and a linear degradation of the batteries (for five years = x 20 / = 18.2%), the battery capacity at the end of the mission will be ( ) x 800 W = 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 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

22 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 = = 25.5 minutes Station keeping energy in dark = 130W x 25.5 / 60 = Wh total energy the system needs to draw from the batteries in the dark side = 60Wh Wh = Wh discharge percentage = Wh / 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 Wh Wh = Wh. The energy available is 620Wh, the energy available to recharge the batteries is 620Wh Wh = Wh. 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 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

23 Figure 11: EyeON heat-environment (Millina F. Diaz-Aguada et al.) 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 Antenna Pointing Sun Sensor Star Sensor Reaction Wheel Power System Flight Computer GPS Receive S-Band Antenna Xenon Propulsion System Page 22

24 X-Band Antenna 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 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

25 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 Solar Constant [W/m²] Albedo Coefficient Earth IR [W/m²] Internal Heat [W] Temperature Limit [K] Area perpen. to Sun [m²] 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

26 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 W. 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 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

27 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 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

28 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 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

29 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 kgm2 ± 5000 rpm 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