Flying Laptop (FLP) is the first minisatellite of the IRS (Institute of Space Systems -Institut für Raumfahrtsysteme) at the University of Stuttgart, Germany. The primary mission objective is to demonstrate and qualify new small-satellite technologies for follow-up missions as summarized in Table 1.
The Flying Laptop project is being developed at IRS. The satellite's subsystems and key functions such as the on-board software and the FDIR (Failure Detection, Isolation and Recovery) concept are allocated to about 10 doctoral candidates. All components, except the ACS (Attitude Control Subsystem ) units, are new developments and mostly engineered in cooperation with industry. While specifications are composed according to mission and project requirements by the students, the satellite component quality benefits from the experience and the industrial procedures that are applied at a professional supplier company. The project progress is observed and supported by the project manager at IRS. In addition, two project advisors (former IRS students) from Astrium GmbH in Friedrichshafen, and one from Tesat-Spacecom GmbH in Backnang, Germany, support the satellite development with their technical knowhow and management advice.
The mission objectives of Flying Laptop include also scientific observations. 2)
• One goal is to use MICS (Multispectral Imaging Camera System) for various scientific Earth observation experiments. Particular targets will be imaged in multispectral bands from different angles in support of BRDF (Bi-directional Reflectance Distribution Function) studies. This in turn implied the introduction of a “spotlight” attitude mode, in which the complete spacecraft is rotated in its orbital path for target fixation. In Flying Laptop notation, the spotlight mode is also referred to as “Target Pointing Mode.” 3) 4)
• Another mission goal, conducted in cooperation with the DTU (Technical University of Denmark), is to utilize the satellite's star trackers for the detection of NEOs (Near Earth Objects). NEOs are asteroids and meteoroids whose trajectories are passing Earth in relatively close orbits.
Figure 1: Artist's rendition of the deployed Flying Laptop spacecraft (image credit: IRS Stuttgart)
Like most satellite projects, the Flying Laptop experienced several reconfigurations throughout its development life. Reshaping the design was necessary due to payload changes, but also due to changing mission requirements . The new configuration is shown in Figure 2 (Ref. 2).
The spacecraft structure is a box of size 60 cm x 70 cm x 80 cm with a total mass of ~ 120 kg and a design life of 2 years. The design is modular. The mechanical structure of the satellite allows a convenient access to the internal components and ensures simultaneous integration of the modules. All segments are made of aluminium due to its high heat conduction properties. The system design is one-failure tolerant to provide secure operations and to preserve system functions. All optical instruments are mounted onto an optical bench, consisting of a CFRP (Carbon Fiber-Reinforced Plastic) sandwich with an aluminium honeycomb core, to insure optical alignment and to minimize thermal expansion effects. Flying Laptop is a low-cost project by using COTS (Commercial-Off-The-Shelf) components whenever possible. 5) 6) 7) 8) 9) 10) 11)
Figure 2: Two views of the Flying Laptop satellite configuration (image credit: IRS Stuttgart)
Table 2: Summary of the spacecraft subsystems and payloads (Ref. 2)
OBC (On-Board Computer): A main requirement is the development of an ultra compact and high-performance OBC, intended to support a RTEMS (Real-Time Executive for Multiprocessor Systems) operating system, a PUS (Packet Utilization Standard) standard based onboard software (OBSW) and a ground/space communication standard based on CCSDS (Consultative Committee for Space Data Systems) protocols. The OBC system consists of four functionally differing boards (Table 3). Each board is available twice for redundancy reasons. All boards are cross-coupled via SpaceWire with a maximum transfer rate of 10 Mbit/s (Ref. 2). 12) 13)
Table 3: Spacecraft OBC composition and properties
The C&DH (Command and Data Handling) functions in the on-board S/W are based on ESA's PUS implementation. PUS describes the transfer of TC and TM data between ground and satellite. In nominal support operations, only one OBC core board and one I/O board are activated controlling the satellite. Since both OBC core boards are cross-coupled to both I/O boards, every link configuration of the boards can be applied for operations. Hence, there are 4 connections available for Flying Laptop to provide failure tolerance against failing boards. To safeguard the accessibility of FLP, two CCSDS boards are available. Both CCSDS boards are operated in parallel and are each connected to one transceiver. As the designation of the boards indicate, the FLP is applying the CCSDS standard as communication protocol for S/C telecommands (TCs) uplinked from ground as well as telemetry (TM) transmitted to ground. All PCBs (Printed Circuit Boards) of the OBC are scaled according to a 3U Eurocard size and mounted into an aluminum frame. The single frames are stacked together and cross-strapped under the front panel (Figure 3). All interfaces that connect to the remaining S/C components are located on top of the stack.
Figure 3: Illustration of the Flying Laptop OBC stack assembly (image credit: IRS Stuttgart)
The OBC boards are SBC (Single Board Computers) designed around the Aeroflex LEON3FT (Fault Tolerant) processor. The LEON3FT is a 32 bit SPARC TM V8 microprocessor with a number of available on chip interfaces including cPCI, SpaceWire and CAN. The OBC boards are implemented with the SpaceWire interface as the primary method of in flight communication. OBC memory resources include 8 MB of on board SRAM and 4 MB of non-volatile memory. Both the SRAM and non-volatile memory interfaces have the LEON3FT on chip EDAC (Error Detection and Correction) designed into the board. The EDAC is capable of detecting two errors on the SRAM or NV memory bus and correcting one. All 4 of the LEON3FT SpaceWire ports have been implemented on the OBC and the data rate is set to 10 Mbit/s. 14)
In addition, ports 3 and 4 support the RMAP (Remote Memory Access Protocol) that gives the user the ability to DMA data directly into SRAM from either port 3 or port 4 SpaceWire interfaces.
Figure 4: Block diagram of the OBC processor board (image credit: Aeroflex, IRS Stuttgart)
EPS (Electrical Power Subsystem): Electrical power is being provided by three solar panels, two of which are deployable (total area of approx. 1 m2). Use of triple-junction GaAs solar cells with an efficiency of 25.3%. In addition, the satellite is being used as an on-orbit testbed for the new generation of penta-junction solar cells with an efficiency higher than 30%. These cells are only used for demonstration (technology evaluation), they are not considered for energy supply and are being mounted on the side panels.
Battery: To achieve a low total cost of the battery system, commercial off-the-shelf Lithium iron phosphate cells manufactured by A123 Systems are used. The system consists of three battery cell strings, yielding a total nominal voltage of 23.1 V and a total nominal capacity of 35 Ah. As the battery cells are susceptible to overcharging, a charge control system to report overvoltages to the PCDU (Power Control and Distribution Unit) is part of the battery system. Furthermore, a simple cell balancer design using bleed resistors is implemented. In order to keep the temperature within the operational temperature range of the cells, the battery system is thermally isolated from the rest of the satellite system by means of multi-layer insulation and glass fiber reinforced plastic parts. 15)
The battery system consists of three battery strings, one for each solar panel. The battery strings for the two deployable solar panels consist of 35 battery cells each, with blocks of five cells in parallel and seven of these blocks in series , yielding a total nominal capacity of 12.5 Ah at a voltage between 18.9 V and 25.2 V. For the string connected to the body mounted solar panel, there are only four cells in parallel, reducing the nominal capacity to 10 Ah, because of the fewer solar cells and the higher temperature of this panel. Thus the whole battery has a nominal capacity of 35 Ah. The mass of the battery system is ~10 kg, its dimensions are 21 x 20 x 20 cm.
Figure 5: Illustration of the battery system (image credit: IRS Stuttgart)
The PCDU (Power Control and Distribution Unit) functions exceed the conventional functional scope of power distribution units. Besides the classic power distribution and regulation functions ,the PCDU serves additionally as the reconfiguration unit for the OBC sub units - processor boards, I/O- and CCSDS boards. The communication between OBC and PCDU is performed via CCs (Common Commands) with a working OBC core board, I/O board and PCDU (Figure 6).
For CC communications, the PCDU provides nominal and redundant cross-coupled, fullduplex communication interfaces in RS422 level (8-N-1) with a baud rate of 115200. The PCDU confirms the proper reception of every command sent by the OBC with a confirmation return. In this way, the OBC supervises power regulation and can request PCDU acquired TM.
The PCDU is being developed in cooperation with an experienced industrial partner (Vectronic Aerospace, Berlin). The device represents the independent monitoring unit for the OBC system and facilitates its operational recovery. Important housekeeping data that is collected at the PCDU is polled by the OBC in a regular interval of 10Hz by a Common Command. If the PCDU is not being polled as specified, a fault of the OBC system is probable.
The PCDU has a size of 220 mm x 160 mm x 118 mm and a mass slightly over 4 kg. Five frame stacks, each corresponding to one PCB, are assembled to a single unit and closed by a cover plate. The PCDU is designed radiation tolerant to at least 20 krad in order to account for the radiation load that is to be expected for a mission life of 2 years. A PCB internal heating for the CPU PCB facilitates the fast warming up to -20ºC in order to prevent damaging of electronic parts due to thermal tension over high temperature gradients. The PCDU is qualified to a lower temperature limit of -40ºC for operational use in order to increase the availability of the PCDU and thus S/C system safety.
Figure 7: Photo of the EM (Engineering Model) PCDU (image credit: Vectronic Aerospace)
Considering the importance of the PCDU for satellite operations a single-point failure tolerant design is particularly realised for C&DH functions inside the unit. Two redundant Central Processing Units (CPUs) are implemented in the PCDU. Both are operated in a hot-redundant concept with a master and a slave unit. The master unit performs all actions, whereas the slave monitors the master. Both CPUs are connected by a toggle logic, which switches the master unit as soon as the currently operating CPU is not responding any more. The master CPU is sending a confirmation signal in a specified period in order to confirm its operability. If this condition is not met, the slave unit commands the toggle-logic to switch the master unit (Figure 8).
Figure 8: Functional design of the CPU toggle-logic (image credit: IRS Stuttgart)
ACS (Attitude Control Subsystem): Flying Laptop is 3-axis stabilized. The requirements call for high-accuracy pointing (150 arcsec or 0.042º) and agile maneuvering capabilities for the imaging mission. The ACS actuators feature four reaction wheels and three magnetic torquers (these are torque rods for momentum dumping of the reaction wheels). Attitude sensing is provided by two 3-axis magnetometers, two coarse sun sensors (6º rms pointing), four fiber-optic rate sensors, one autonomous star tracker (fine pointing accuracy of < 2 arcsec), and three GPS receivers (GENIUS). 17) 18) 19)
Table 4: Pointing parameters of the ACS
Following is a description of the various ACS and S/C subsystem components:
• In this context, GENIUS (GPS Enhanced NavIgation system for the University of Stuttgart micro-satellite) is an onboard experiment being conducted in cooperation with DLR/GSOC (Figures 9 and 10). GENIUS consists of three COTS Phoenix GPS boards. Each of the receivers is connected to separate GPS antenna via a low noise amplifier. The antennas of three separate GPS receivers are being placed on three corners of the body-mounted central solar array in an L-shape configuration. The GENIUS performance offers real-time position, velocity and timing information with estimated accuracies of 10 m, 0.1 m/s and 1 µs, respectively (in addition attitude is being provided). The Phoenix GPS receiver is a commercial GPS receiver board with a new DLR/GSOC developed firmware for space and high dynamics applications. The receiver has 12 tracking channels and is able to measure phase and Doppler shift of the GPS-L1 carrier signal.
Figure 9: Configuration of the GENIUS system (image credit: DLR/GSOC)
Figure 10: Illustration of GPS antenna allocations (image credit: IRS)
The GENIUS GPS system consists of three independent GPS receiver boards, each connected to a separate antenna and low noise amplifier (LNA) as shown in Figure 9. The GPS Box is connected to the on-board computer (OBC), the power control and distribution unit (PCDU) and the ultra stable oscillator (USO). The used Phoenix boards are commercial 12-channel GPS L1 receivers with a DLR/GSOC developed firmware for space and high dynamics applications. Three GPS antennas are mounted on the middle solar panel in an L-shaped arrangement, creating two baselines with a length of 440 mm and 610 mm respectively (Figure 10). The three GPS receivers are integrated in a single 100 mm x 80 mm x 67 mm box together with an interface board for RS-422 conversion, the USO signal distributor and a latch-up protection for each receiver. To achieve a high level of redundancy, each receiver can be switched on/off independently varying the system input power from 0.9 W for 1 receiver to 2.6 W for all 3 receivers according to measurements at the testing model.
An algorithm based on a Kalman Filter is used to process the measurement data and produce an offline attitude solution which will be compared to the attitude information available from the satellite's star camera. The algorithm uses the lambda-method to resolve the integer ambiguities of the double differences of the carrier phase measurements. These resolved double difference ambiguities are then used to fix the single difference ambiguities in the filter. Hence, the algorithm provides a seamless transition from the ambiguity resolution to the attitude determination.
• STR (Star Tracker): The autonomous star tracker in the ACS configuration is the newly developed µASC (micro Advanced Stellar Compass) of DTU (Technical University of Denmark), Lyngby, Denmark. In fact, µASC is of ASC heritage flown on Orsted, SAC-C, CHAMP, GRACE, ADEOS-2, SMART-1, GOCE, etc. The µASC instrument is physically divided into a µDPU (micro Data Processing Unit) with hot/cold redundancy and CHU (Camera Head Unit), a µDPU may drive up to 4 CHUs (2 CHUs are being used on Flying Laptop). The intrinsic accuracy of an attitude measurement from a single CHU is better than 1 arcsec at an integration time of 0.5 s. This attitude is autonomously calculated based on all brighter stars in the FOV of the CHU. The µASC on Flying Laptop provides a pointing knowledge within 2 arcsec. Furthermore, µASC delivers attitude information at S/C angular rates of up to 10º/s and thus enables rapid repointing of the platform to any object. This attitude is autonomously calculated based on all brighter stars in the FOV of the CHU. The µASC on Flying Laptop needs to provide a pointing knowledge of one pixel at 7.4 arcsec.
The use of µASC on Flying Laptop is an early spaceborne demonstration of this instrument. Currently PROBA-2 is another mission under development using the µASC device (launch planned for late 2007). 20)
Figure 11: Illustration of the µASC instrument (image credit: IRS)
Figure 12: View of a camera head unit and baffle (image credit: IRS)
• The three magnetic torque rods employed are developed by ZARM Technik, Bremen, with a linear dipole moment of 6 Am2. The torquers are connected to a power box that includes two I2C buses for connection to the OBC. The whole system is single redundant.
• Magnetometer: ZARM Technik provides also the AMR (Anisotropic-Magneto-Resistive) magnetometer, a microcontroller-based 3-axis magnetometer with digital output. Two magnetometers are being installed on the microsatellite. The Earth's magnetic vector field is being used as input information for the magnetic torquers (detumbling after launcher separation, etc.). The ARM sensor is the HMC-1023 model of Honeywell.
• The angular rate of the S/C is measured with 4 single-axis COTS fiber optic rate gyros (FOGs) in a tetrahedron configuration. The sensors employed are C-FORS (Commercial Fiber Optic Rate Sensor) of Litef. The complete FOG assembly has a mass of ~1.7 kg.
Figure 13: Functional architecture of the spacecraft (image credit: IRS Stuttgart)
Figure 14: Overview of the ACS sensors and actuators (image credit: IRS Stuttgart)
RF communications: For telemetry and telecommand, S-band (low and high gain) antennas are being installed on the satellite. The S-band command uplink has a frequency at 2.068 GHz, a telemetry downlink at 2.245 GHz, and a data downlink at 2.425 GHz.
Flying Laptop also features commanding by so-called HPCs (High Priority Commands) of the PCDU in case of an emergency, indicated by the red colored communication path in Figure 15.
Figure 15: Block diagram of the communications system (image credit: IRS Stuttgart)
The satellite's Cassegrain system with its 50 cm primary dish provides the antenna reflector for the Ka-band communication and is also used as the optical system element for the thermal infrared camera (TICS). The TWT is the design driver for the Li-ion battery system (50 Ah, 6 cells) to handle its high power requirement (for a maximum duration of 20 minutes of operations support).
The microsatellite will be operated by students at IRS (Institut für Raumfahrtsysteme), of the University of Stuttgart. The existing ground station on campus is being upgraded to permit satellite communications in the following frequency bands: UHF, VHF, L-band, S-band and Ka-band (payload support).
Figure 16: Block diagram of TM/TC FPGA device (image credit: IRS Stuttgart)
Launch: The launch of the Flying Laptop satellite as a secondary payload is planned for late 2014 on a PSLV launcher from India.
Orbit: A sun-synchronous polar circular orbit with an altitude of about 700 km.
Sensor complement: (MICS, TICS, AIS)
The scientific payload of the satellite is a triple imaging system, a VNIR (Visible Near-Infrared) system called MICS, a TIR (Thermal Infrared ) camera, and a Ka-band communication/imaging system (called TICS). The last two instruments are intended to make dual use of a cassegrain mirror system.
MICS (Multispectral Imaging Camera System):
The objective is to observe in the VNIR range of the spectrum in three bands at medium resolution (GSD of 25 m). MICS consists of three single cameras, each with an area array CCD detector for snapshot observations.
Table 5: Key parameters of the MICS instrument
The optical system uses a double Gauss telescope with interference filters placed in front of the system. The use of an area array detector has advantages for the measurement of the BRDF and allows easier referencing on ground. The BRDF measurements will be done in the target-pointing mode, where the satellite is focused on the target site during the whole passage.
In order to accomplish reliable scientific measurements, periodic calibration of the instrument is mandatory, not only on ground, but also in space. A particular LED (Light Emitting Diode) device is being used to verify the following items:
• Pixel-to-pixel shift (flat-field calibration)
• Spectral shift of interference filters
• Radiometric performance.
Figure 17: Schematic sectional drawing of MICS (image credit: IRS Stuttgart))
TICS (Thermal Infrared Camera System):
The objective is to observe in the TIR (Thermal Infrared) wavelength ranges of 8.2-9 µm and 10.6-12.4 µm, using an uncooled microbolometer detector array of size 320 x 240 pixels (snapshot imagery). The detector is temperature stabilized and cooled by Peltier elements in order to achieve the desired SNR (Signal-to-Noise Ratio).
The TICS optics system consists of a Cassegrain system (f/1.6), designed as a dual-band system (it serves as a telescope for the TIR range, and as antenna for the Ka-band signals), as well as relay optics. The primary mirror (500 mm aperture diameter) of the Cassegrain optics system as well as the retaining structure of the secondary mirror are being produced from CFRP (for temperature stability and alignment). The TICS instrument is providing a GSD (Ground Sample Distance) of 100 m in TIR.
Pointing modes for image acquisition:
Three attitude control modes have been defined for image acquisition:
1) Inertial pointing mode: In this mode, the star sensor is being used to provide high accuracy pointing knowledge of 7.5 arcseconds. The satellite will be and will also remain inertially stabilized, i.e. the coordinate system of the satellite will maintain in the same orientation with respect to stars. Note: This mode is not useful for taking Earth imagery. Other observations (e.g. stars, moon) are possible.
2) Nadir pointing mode: In this mode the satellite is being aligned in the direction of the Earth, i.e. the z-vector of the satellite's coordinate system is perpendicular to the Earth's surface; hence, the angular rate remains constant. This mode is also called the “Earth-pointing mode,” being used for image acquisition, attenuation measurements and trace gas detection.
3) Target-pointing mode (or spotlight mode): This mode is being used to achieve the required coverage/resolution of the planned scientific observations. In this mode, the S/C points at fixed target (spot an Earth's surface) during an extended period of time thereby achieving a TDI (Time Delay Integration) effect. The spotlight service requires a slewing of the S/C to keep the instruments pointed. The maximum slew rate for this maneuver is 1º/s. This is the most demanding support mode of the satellite in terms of control algorithms.
Figure 18: Imaging modes: A) Inertial-, B) Nadir- and C) Target-pointing mode (image credit: IRS Stuttgart)
Research experiments and technology demonstrations:
The sensor complement and the Ka-band antenna are intended to be used as research tools in the field of Earth observation through remote sensing. The following topics will be investigated:
• BRDF (Bi-directional Reflectance Distribution Function) measurements: In the target pointing mode, this function is measured in different spectral bands (visible, near infrared, and thermal infrared). The cameras take imagery continuously of the same target area during a small segment of the orbit - at various observation angles, resulting in bi-directional reflectance measurements. Obviously, homogeneous ground surfaces like large forests or deserts represent ideal targets for this support mode.
BRDF specifies the behavior of surface scattering as a function of illumination and view angles at a particular wavelength. BRDF is defined as being the ratio of the reflected radiance to the incident flux per unit area. As such, the BRDF function plays a decisive role in the analysis of spaceborne remote sensing data.
• Demonstration of precipitation measurements in Ka-band: Experiments have shown, that the differential radio signal attenuation in a horizontal path through rain in two different frequency ranges (between 10 and 40 GHz) is linearly dependent on the rain rate. The target pointing mode is being used to acquire precipitation measurements in Ka-band and Ku-band. Note: The rainfall determination requires the differential measurement of the signal attenuation at two two distinct frequencies. For the Ka-band, the second frequency should be in the Ku-band (this is being implemented by another onboard Ku-band transmitter).
The Ka-band antenna with its large bandwidth is being used for the study of atmospheric attenuation within this frequency range. The Ka-band signal is influenced by a variety of causes in the atmosphere:
- Attenuation through rain
- Attenuation through clouds
- Attenuation through atmospheric constituents
- Instability of the atmospheric refraction index
- Phase transformations from ice particles to water droplets
The objective is to retrieve the total content of various trace gases.
• Multispectral observations: Simultaneous observations are planned with all cameras. In addition, the Ka-band antenna with its high power will also be used as a radar transmitter (providing a GSD of 25 km). The signal will be transmitted from space, but the reflected signal needs to be captured with the help of measurement towers on the ground, not at the satellite. The broadband Ka-band signal permits data rates of up to 100 Mbit/s.
• The implementation of the TWT amplifier with a transmission power of 57 W represents a new capability for microsatellite operations.
• FPGAs: The introduction and reliance of onboard computing with an FPGA system represents a new approach to conventional system architectures. It provides the capability to directly generate the logical configuration of FPGA gates from a C-like high level language without producing the machine code for a processor (hence, massive parallel processing is possible). Using an onboard computer architecture with several reciprocative checking FPGAs, a safe system is obtained that even exceeds the performance of current PCs through its ability of parallel real-time processing. An inherent advantage of FPGA architectures is the capability of reconfiguration within milliseconds.
To make the system fault-tolerant and to address radiation issues, four equal independent nodes work together. Depending on the state of the system, 1-4 nodes may run in parallel; they may be switched on or off dynamically.
• NEA (Near Earth Asteroid) detection. Aside from delivering attitude information, the star tracker in use possesses a built-in feature to automatically detect, identify and track any other faint luminous object, not being a star, as long as the object is brighter than the visual (or apparent) magnitude Mv 11.
The Technical University of Denmark (DTU) has proposed an interplanetary mission to search for Near Earth Asteroids (NEA), based on their star tracker, µASC. Observation time for these science experiments of the Flying Laptop will be made available to test and verify this concept, in the eclipse phase of the LEO orbit (for further information see reference 6).
• Rent-a-Sat mode: The inherent high flexibility of the onboard computer system will be used to operate the Flying Laptop in a so-called 'Rent-a-Sat' mode. Interested parties (companies or institutions) can rent the satellite as a development platform in space. It is possible to configure the system for customer preferences (i.e., the characteristics of a certain processor can be simulated through the hardware). With this versatility the 'Rent-a-Sat' system is well suited for spaceborne software or firmware validations.
• PanCam (Panoramic Camera) is an additional COTS camera on Flying Laptop to provide context color video imagery of Earth . PanCam is required because the narrow FOVs of the two science imagers (MICS and TICS) are insufficient for context information imagery to increase public outreach of the Small Satellite Program. PanCam uses CMOS technology with a pixel pitch of 6.7 µm. It has 1280 x 1024 pixels and can capture up to 27 images per second in full resolution. Using a focal length of approximately 25 mm, the sensor can cover a FOV of 20º x 16º. This results in a swath width of approximately 250 km and a ground sample distance of around 200 m from a 700 km orbit. The video link of PanCam employs a lossy video compression technique to be able to handle the large source data volume. 21)
AIS (Automatic Identification System):
In Q3 of 2012, an AIS receiver inclusive antenna was implemented as a new payload on Flying Laptop. This payload was developed, build and tested by the DLR Institute of Space System in Bremen. The objective of the AIS instrumentation is to receive AIS signals from ships in the ground segment. 22)
Since January 1, 2004, it is mandantory to run an AIS transmitter for all ships bigger than 300 GRT in international waters. Since Juli 1, 2008, all ships in national waters bigger than 500 GRT also need to run and AIS transmitter. AIS, is a system to supervise marine traffic. In times of increasing ship traffic, a system like this is indispensable. AIS shall be used or the following:
- preventing collisions
- information for adjoining coastal states considering ships and their cargo
- appliance for landward survilliance.
The system works as follows: Ships will send a message in regular time intervals. These messages contain, among other information the position, route, and velocity of the ship, the ship name and the call sign. If a ship has an AIS receiver on board, it can use the signals for better planning and desicion making.
The AIS signals can also be received from a spacecraft. In a last-minute cooperation, between DLR and the IRS (Institut of Space Systems) of the University of Stuttgart, an AIS receiver incl. antenna could be accomodated within the satellite. According to the slogan, form follow function, the unusual form of the AIS receiver housing was created.
Figure 19: Photo of the AIS instrumentation (image credit: IRS Stuttgart)
The ground segment of the Flying Laptop mission of IRS (Institute of Space Systems) at the University of Stuttgart is being setup as a collaborative network of several stations on three continents to extend the mission control capabilities and the data return. The overall intent of the collaboration is to support small satellite missions and to contribute to a global partnership based network. 23)
A ground station network collaboration was formed between CASPER (Center for Astrophysics, Space Physics & Engineering Research) at Baylor University of Waco, TX, USA and IRS of Stuttgart University, Germany. Baylor University is building a new ground station with facilities at BRIC (Baylor Research and Innovation Collaborative), a former factory complex that was recently rebuilt. Thus, efforts were made to plan the ground station in a way that supports its functionality in an optimum manner, beginning with the allocation of most suitable rooms in the early renovation design process. Since there is not yet an own satellite mission established or planned at CASPER, the focus of the ground station’s architecture is set on the flexibility to support various satellite missions.
The first mission support of the CASPER ground station will be Flying Laptop of IRS. Furthermore, CASPER intends to support microsatellite missions that are initiated by the SRL (Space Robotics Laboratory) at Tohoku University, Sendai, Japan. These three universities will connect their ground stations to a network, which is characterized by optimal compatibility, close collaboration and mutual trust.
Table 6: Satellites from collaborating institutions
First steps in the design of a satellite ground station at CASPER were made. The requirements are defined, frequency ranges determined and link budget analyses performed. Investigations on the S-band antenna were made, referring to solutions on the market and the installation on the roof of the BRIC. Furthermore, the functional design of the Mission Control Center and related rooms is finished.
1) Information provided by Jens Eickhoff, Astrium GmbH, Friedrichshafen, Germany
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7) “Der Kleinsatellit Flying Laptop,” 2006, URL: http://www.tz-raumfahrt.de/pdf/FLP_Projektbeschreibung.pdf
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9) H.-P. Roeser, F. Huber, G. Grillmayer, M. Lengowski, S. Walz, A. Falke, T. Wegmann, “A small satellite of the University Stuttgart - a demonstrator for new techniques,” Proceedings of the 31st International Symposium on Remote Sensing of Environment (ISRSE) at NIERSC (Nansen International Environmental and Remote Sensing Center), Saint Petersburg, Russia, June 20-24, 2005
10) T. Kuwahara, F. Huber, A. Falke, M. Lengowski, S. Walz, G. Grillmayer, H.-P. Röser, “System Design of the Small Satellite Flying Laptop, as the Technology Demonstrator of the FPGA-based on-board Computing System,” 58th IAC (International Astronautical Congress), International Space Expo, Hyderabad, India, Sept. 24-28, 2007, IAC-07- B4.6.08
11) Fabian Steinmetz, Michael Lengowski, Daniel Winter, Lucas Salvador, Hans-Peter Röser, Pierre Rochus, “Validation of the Structural-Thermal-Model of the Small Earth Observation Satellite Flying Laptop,” Proceedings of the 9th IAA Symposium on Small Satellites for Earth Observation, Berlin, Germany, April 8-12, 2013
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13) Jens Eickhoff, Barry Cook, Paul Walker, Sandi Habinc, Rouven Witt, Hans-Peter Röser, “Common Board Design for the OBC I/O Unit and The OBC CCSDS Unit of The Stuttgart University Satellite Flying Laptop," Proceedings of the DASIA (DAta Systems In Aerospace) 2011 Conference, San Anton, Malta, May 17-20, 2011, ESA SP-694, August 2011
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16) Alexander N. Uryu, Rouven Witt, Michael Fritz, Samson Houssou, Jens Eickhoff, Hans-Peter Röser, “Multifunctional Power Control and Distribution Unit for Command Chain Reconfiguration,” Proceedings of the 4S (Small Satellites Systems and Services) Symposium, Portoroz, Slovenia, June 4-8, 2012
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.