Minimize Flying Laptop

Flying Laptop

Spacecraft    Launch    Sensor Complement    Ground Segment    References

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.

Technology introduction

Functions to be demonstrated

UT 699 LEON 3 FT microprocessor

High-grade main OBC (On-Board Computer)

SpaceWire application on university small satellites

Communication between core control units via SpaceWire

CCSDS protocol for communications

Standardized RF communication protocol

New implementation design for recovery of S/C control functions through PCDU (Power Control and Distribution Unit)

- Reconstitution procedure of LEON3FT OBC components
- S/C control by high priority commands

Lithium iron phosphate (LiFePO4) battery

Use as secondary power source

High-efficiency solar cell test string

Part of primary power source

ACS (Attitude Control Subsystem) components for a pointing performance ≤ 150 arcsec

Highly accurate ACS for communication and multiangular target pointing by adjusting the complete satellite

New implementation method for a solar panel release mechanism

Deployment of side solar panels

Payload OBC

Support of payloads with FPGA computing capability

Experimental laser link system OSIRIS

High-speed optical communications

Table 1: Overview of Flying Laptop technology demonstrations 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 Airbus DS 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) 5)

• A further scientific 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.

• A further scientific mission goal is the observation of ship traffic. By means of an AIS Receiver in combination with the optical observation of MICS the spaceborne ship surveillance shall be evaluated.


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 90 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. The lower part consists of integral aluminum parts and therefore offers a cost-effective and precise assembly, adjusted force transmission and good thermal properties. The upper part, where the optical systems are installed, consists of carbon‐fibre reinforced sandwich structures. These offer a low mass, high stiffness as well as low thermal longitudinal expansion to provide a stable calibrated alignment of the cameras in the satellite and to each other. 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. 6) 7) 8) 9) 10) 11) 12)


Figure 2: Two views of the Flying Laptop satellite configuration (image credit: IRS Stuttgart)




C&DH (Command & Data Handling)

OBC (On-Board Computer)

Monitoring and control of the spacecraft

EPS (Electrical Power Subsystem)

PCDU (Power Control & Distribution Unit)

- Power regulation and distribution
- Command Chain Reconfiguration
- Collection of analogue sensor data

Solar panels

Primary energy source – GaAs solar cells with an efficiency of 25.3% for a max power generation of 269.2W with 3 solar panels at 1 m2 surface area


Secondary energy source: Three LiFePO4 cell strings with a capacity of 35 Ah for high power operations and eclipse phase, nominal voltage of 23.1 V.


Data and power connections between on-board components

ACS (Attitude and Control Subsystem) for 3-axis stabilization

Reaction wheels

Main actuator to control satellite's attitude by changing the momentum of spinning wheels, 4 units in tetrahedron configuration


Secondary actuator for emergency situations and LEOP, also used for wheel desaturation, redundant electro-magnetic coil for each axis which is producing a control torque by interacting with Earth's magnetic field

Star Trackers

Measurement of attitude with the help of star constellations, two camera head units to avoid blinding

FOGs (Fiber Optic Gyros)

Measurement of rotational rate, 4 units in tetrahedron configuration

Coarse sun sensors

Acquisition of sun direction, at least two redundant units for each axis

GPS receiver

Acquisition of position and speed of satellite, three redundant units

Spacecraft structure


Supporting chassis (box structure) for on-board components
- Size: 60 cm x 70 cm x 90 cm
- Mass: ~120 kg

Retaining/deployment mechanism

Mechanism to fix the side solar panels during transport and launch. The fuse wires, which fasten the deployment mechanism for the side solar panels, are melted in orbit

De-orbit mechanism

Deployable foil with an maximum expanse of 2.5 m x 2.5m to facilitate de-orbiting within 25 years

TCS (Thermal Control Subsystem)

SSM Radiator

Radiation of dissipated energy to space


Insulation cover of the satellite protecting form temperature differences of space

RF communications

Transceiver system

- Consists of Telecommand (TC) receiver and Telemetry (TM) transmitter, both are operated in S-band
- Payload S-band downlink up to 10 Mbit/s (with horn antenna)


Optical communications

OSIRIS (Optical high-Speed Infrared Link System)

Laser link terminal for optical data transmission up to 100 Mbit/s at 1550 nm

Payload OBC


Monitoring and control of the payloads and data handling with SRAM FPGA


Multispectral Imaging Camera System

Earth Observation in three spectral bands:
Green: 530 – 580 nm
Red: 620 – 670 nm
NIR: 835 – 885 nm



Panoramic image acquisition for orientation purposes

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). 13) 14)

OBC core board

- features a UT699 LEON3 with a fault tolerant 32 bit SPARC V8 microprocessor from Aeroflex Gaisler
- is operated with a clock speed of 33 MHz
- SpaceWire is utilized as communication method
- qualified for space use from -40ºC to +105ºC and radiation-hardened up to 300 krad
- an error correction and detection functionality protects the OBC memory from faults

I/O board

- is operated by a radiation-tolerant flash FPGA with non-volatile memory, implemented by 4Links
- incorporates all digital interfaces to the satellite components, except the payload components


- similar technical composition as the I/O board, but less digital interfaces
- acts as the TC/TM pre-processor board:
a) Decoder of commands from transceiver/ground; b) Encoder for housekeeping data to transceiver/ground
- forwards Common Commands (CC) to OBC core board


- a reconfiguration unit being embedded in the PCDU (Power Control and Distribution Unit)

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. 15) 16)

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 atest string of a new generation of 100 µm 3G triple junction solar cells.

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. 17)

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.


Figure 6: Overview of Interface Connections between OBC System and PCDU (image credit: IRS Stuttgart) 18)

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, eight 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). 19) 20) 21)

Pointing knowledge, absolute

±7 arcsec (±1 pixel)

Pointing knowledge, relative

±2.5 arcsec (±1/3 pixel)

Pointing accuracy

±150 arcsec (± 20 pixel)

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 microsatellite) 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) and the power control and distribution unit (PCDU). 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. 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 Ørsted, 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. 22)


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 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, and S-band.


Figure 16: Block diagram of TM/TC FPGA device (image credit: IRS Stuttgart)


Figure 17: Sensor positions during vibration test (image credit: IRS Stuttgart)


Project development status:

• March 9, 2017: The Flying Laptop passed its Flight Readiness Review as well as its Operational Readiness Review on Tuesday, March 7, 2017! The review board, consisting of representatives from DLR, Airbus, Tesat and Thales, confirmed the success of the reviews, provided that the actions for launch and operation preparation are completed and documented as planned before the satellite is shipped, and therefore also before its launch. 23) 24)


Figure 18: Photo of the Flying Laptop Team in 2017 (image credit: IRS Stuttgart)

• February 2017: The Flying Laptop minisatellite is complete and tested. Delivery to the launch site should occur in April (Ref. 30). 25)

• As of the summer of 2015, the flight software of the Flying Laptop is being finalized and the last functional system tests will be prepared and executed. Thus, by the end of 2015 the satellite will be completed.

• Since mid-2014, the fully integrated small satellite Flying Laptop is in the system testing phase at the Institute of Space Systems located at the University of Stuttgart in Germany.


Launch: The Flying Laptop satellite was launched on July 14, 2017 (06:36:49 UTC, 09:36:49 Moscow time) as a secondary payload on a Soyuz -2.1a Fregat-M vehicle configuration from Baikonur, Kazakhstan. The primary payload on this flight was the Kanopus-V-IK (IK =Infra-Krasny, means "infrared" in Russian) mission of Roscosmos. 26) 27) 28) 29) 30)

Orbit: A sun-synchronous polar circular orbit with an altitude of 600 km, inclination = 97.6º.

Secondary payloads: In total, 72 secondary satellites will be launched on the Kanopus-V-IK mission, including spacecraft for four separate commercial remote sensing and weather constellations.

Glavkosmos, a subsidiary of Russian state space corporation Roscosmos, is seeking to become a larger provider of rideshare launch services for small satellites. In addition to this upcoming launch, Glavkosmos is planning to fly about 40 more small satellites on two Soyuz missions from the new Vostochny Cosmodrome in Russia's Far East region late this year, with additional launch opportunities planned for 2018 and beyond.

In addition, the German Orbital Systems company of Berlin, ECM Space Technologies GmbH, actively participates in the launch, by supplying a DCSM (Deployment Control and Separation Sequence Management) unit. Integration of two secondary satellites built in Germany, three from Russia and 12 from the USA. GlavKosmos subcontracted this task to ECM. 31)

Forty-eight (48) Dove satellites (Flock 2k) of Planet will be launched into an SSO of 475 km altitude. 32) Planet of San Francisco is the biggest single customer for the upcoming launch. This "flock" of 48 satellites will go into a sun-synchronous orbit, but one slightly different from those of 88 similar satellites launched on an Indian Polar Satellite Launch Vehicle in February. Those satellites went into a sun-synchronous orbit that crosses the equator at 9:30 hours local time. The new satellites will launch into an orbit that crosses the equator at about 11:00 hours local time.

• Spire Global of San Francisco, which is deploying a fleet of CubeSats to collect GPS radio occultation and ship tracking data, has eight (8) of its Lemur CubeSats on this launch.

• GeoOptics of Pasadena, CA, is flying its three CICERO 6U CubeSats (CICERO-1, -2,- 3), designed to collect GPS radio occultation data to support weather forecasting.

• Flying Laptop, a minisatellite (120 kg) of IRS (Institute of Space Systems), University of Stuttgart, Germany.

• WNISAT- 1R (Weather News Inc. Satellite 1R), a microsatellite (43 kg) for north arctic routes and atmosphere monitoring. The project is started from the commercial objects between Weathernews and Axelspace, Japan. The object of the WNISAT-1R mission is monitoring of the Northern sea routes and of the CO2 content of the atmosphere. 33)

• NorSat-1 and NorSat-2 of the Norwegian Space Center, built by UTIAS/SFL (Canada) on GNB (Generic Nanosatellite Bus), each with a mass of 15 kg. The NORSAT-2 satellite will carry a next generation Automatic Identification System (AIS) receiver from Kongsberg Seatex, along with a VDE (VHF Data Exchange) payload that will enable two-way communication at higher data rates than possible with AIS. NorSat-1 (< 30 kg) carries three instruments: An AIS receiver, a Langmuir Probe Instrument, and CLARA (Compact Lightweight Absolute Radiometer), intended to observe total solar irradiation and variations over time.

• TechnoSat is a nanosatellite project (18 kg) of the Technical University of Berlin. Test of new nanosatellite components, including a camera, a new reaction wheel system, a star tracker, a transmitter, a fluid dynamic actuator, and commercial laser retro-reflectors.

• Corvus-BC 1 and Corvus-BC 2: Astro Digital, formerly known as Aquila Space, (formed from the core team of the former Dauria Space daughter Canopus Systems) developed the Corvus-BC Earth observation CubeSats (6U, 10 kg). The Landmapper-BC spacecraft, also known as Corvus-BC1 and Corvus-BC2, have color and infrared cameras for wide-area imaging.

• MKA-N 1, 2: Two Earth observation CubeSats (6U, 10 kg each) of Roskosmos, built by Dauria Aerospace.

• NanoACE, a technology demonstrator by Tyvak Nanosatellite Systems, Inc. (Irvine, CA) to validate the Endeavor suite technologies that will be used for future missions and is solely for the purpose of internal Tyvak development as an attitude control experiment. A 3U CubeSat (5.2 kg).

• Mayak: A Russian 3U CubeSat (4kg) developed by a group of young scientists named "Your sector of space" with support of the Moscow State University of Mechanical Engineering (MSUME). In orbit, the 3U CubeSat will deploy four triangular reflectors, 4 m2 each, which form a tetrahedral shape. The reflectors are made from metalized membrane with reflection coefficient of 95%. The reflector will provide a -10 optical magnitude at the beginning of the flight to allow for easy tracking. Mayak will be put into a tumbling motion over all axes, with at least 1 revolution per second.

• Iskra-MAI-85, a CubeSat of the Moscow Aviation Institute. 34)

• Ecuador-UTE-YuZGU, a CubeSat of the Kursk South-Western State University, Russia.

09:36:49 – launch vehicle lift-off;
09:38:46 – 1st stage separation;
09:41:36 – 2nd stage separation;
09:41:38 – fairing jettison;
09:45:37 – head module separation;
09:45:42 – 09:52:18 – Fregat upper stage flight to a transfer orbit;
10:35:01 – 10:36:27 – Fregat upper stage flight to the Kanopus-V-IK separation orbit;
10:38:07 – Kanopus-V-IK separation (orbit i=97.44°; H = 522.5km; h = 478.6km);
11:13:29 – 11:14:35 – Fregat upper stage flight to the second transfer orbit;
11:58:29 – 11:59:35 – Fregat upper stage flight to the separation orbit of a group of smallsats;
12:01:43 – 12:05:03 – Phase 1. Separation of 5 smallsats (orbits i=97.61°; H = 601.5-600.1km; h = 600.0-590.1km);
12:10:03 – 12:26:43 – Phase 2. Separation of 19 smallsats (orbits i=97.62-97.61°; H = 601.0-606.9km; h = 580.1-587.4km);
12:51:49 – 12:53:15 – Fregat upper stage flight to the third transfer orbit;
13:34:39 – 13:35:51 – Fregat upper stage flight to the separation orbit of a group of smallsats;
17:18:23 – 17:41:17 – Separation of 48 smallsats (orbits i=97.00-97.01°; H = 485.0-477.4km; h = 482.2-450.5km);
17:51:49 – 17:53:45 – Fregat upper stage flight to reentry orbit;
~18:18:49 – Fregat upper stage reentry (altitude – 100 km), sinking in the Indian Ocean.

Table 5: Overview of the Kanopus-V-IK timeline (Moscow time, Ref. 26)



Sensor complement: (MICS, AIS)

The scientific payload of the satellite is a triple imaging system, a VNIR (Visible Near-Infrared) system called MICS, the PAMCAM, and AIS Receiver, as well as the star tracker of the Attitude Control System of the Flying Laptop.

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 20 m). MICS consists of three single cameras, each with an area array CCD detector for snapshot observations.

Spectral bands

530 nm - 580 nm (green)
620 nm - 670 nm (red)
820 nm - 870 nm (NIR)

GSD (Ground Sample Distance)

20 m

Swath width

20 km

Data quantization

12 bit

Instrument mass, power ,size

~4 kg, ~5 W, 100 mm x 90 mm x 400 mm

Table 6: 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 multi-angular measurements. The multi-angular 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 19: Schematic sectional drawing of MICS (image credit: IRS Stuttgart)


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 20: Imaging modes: A) Inertial-, B) Nadir- and C) Target-pointing mode (image credit: IRS Stuttgart)


Research experiments and technology demonstrations:

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, a cold redundant computer system is used.

NEO (Near Earth Object) 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).

PamCam (Panoramic Camera) is an additional COTS camera on Flying Laptop to provide context color video imagery of Earth . PamCam 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. PamCam 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 PamCam employs a lossy video compression technique to be able to handle the large source data volume. 35)


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. 36)

Since January 1, 2004, it is mandatory to run an AIS transmitter for all ships bigger than 300 GRT in international waters. Since July 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 21: Photo of the AIS instrumentation (image credit: IRS Stuttgart)



Ground segment:

To communicate with the Institute's own satellite, the IRS has its own ground station. The objective of the installation of the system is to automate operations as much as possible. Thus, the usage time of the satellite is to be maximized without providing an expensive and complex 24‐hour manual operation. The main tasks of this station are sending commands to the satellite (telecommanding, TC), the reception of Housekeeping data (telemetry, TM), as well as the download of the scientific payload data from the satellite [downlink payload, DDS (Data Downlink System)].

The antenna system with a reflector diameter of 2.5 m is capable to transmit within the commercial space research S‐band (2075 MHz ‐ 2090 MHz) with an equivalent output power of up to 20 kW and simultaneously, also within the commercial space research S‐band (2257 MHz ‐ 2278 MHz) to receive telemetry. Both are done in the right‐hand circular polarization oriented (RHCP). Simultaneously, the antenna system is able to receive the payload data in the left‐hand circular polarization LHCP with up to 10 Mbit/s in the amateur radio S‐band (2400 MHz ‐ 2428 MHz). The transceiver is the commercial satellite transceiver CORTEX CRT of Zodiac Aerospace.

In order to increase the contact time and to auxiliary ground stations in the fault case of the IRS station, a network of ground stations was established.

In LEOP (Launch and Early Orbit Phase), three DLR Ground are being used, all of which allow TM downlink and TC uplink on a live data ling using SLE through GSOC (German Space Operations Center):

• Weilheim (WHM) station in Bavaria, Germany

• Inuvik (INU) station in the Northwest Territories, Canada

• O'Higgins (OHG) station in Antarctica.

The Weilheim ground station shall still be used during the commissioning phase. The IRS ground station shall be taken into operation during this phase and shall support more and more passes. At the end of commissioning the IRS ground station shall be the main operational ground and Weilheim shall only be used in emergencies as well. Also the ground station of GFZ (German Research Center for Geosciences) in Ny Alesund, Svalbard, Norway (NYA) shall be started to be used for automated TM downlink during this phase. This ground station only provides offline TM reception and no uplink.

For the data downlink, two additional ham radio stations at the University Putra Malaysia in Kuala Lumpur (MAL) and a private one in Kiel, Germany (KIE) may be used.


1) Information provided by Jens Eickhoff, Astrium GmbH, Friedrichshafen, Germany

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8) "Der Kleinsatellit Flying Laptop," 2006, URL:

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12) 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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14) 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

15) Sandi Habinc, Barry Cook, Paul Walker, Jens Eickhoff, Rouven Witt, Hans-Peter Röser, "Using FPGAs and a LEON3FT Processor to Build a Flying Laptop," Respace/MAPLD (Military and Aerospace Applications of Programmable Devices and Technologies) 2011 Conference, Albuquerque, NM, USA, Aug. 22-25, 2011, URL:

16) Jens Eickhoff, Kai Klemich, Ulrich Mohr, Nico Bucher, Rouven Witt, Bastian Baetz, Gianluca Cerrone, Wolfgang Heinen, "Operating the Stuttgart Micro Satellite based on the Combined Data and Power Management Infrastructure," SpaceOps 2014, 13th International Conference on Space Operations, Pasadena, CA, USA, May 5-9, 2014, paper: AIAA 2014-1730, URL:

17) Kai-Sören Klemich, Michael Lengowski, Hans-Peter Röser, Justus Speichermann, "A robust and cost-effective approach for a battery system based on off-the-shelf Lithium iron phosphate cells for the small satellite Flying Laptop," Proceedings of the 9th IAA Symposium on Small Satellites for Earth Observation, Berlin, Germany, April 8-12, 2013

18) 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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20) A. Hauschild, G. Grillmayer, O. Montenbruck, M. Markgraf, P. Vörsmann, "GPS Based Attitude Determination for the Flying Laptop Satellite," Proceedings of the 6th IAA Symposium on Small Satellites for Earth Observation, Berlin, Germany, April 23 - 26, 2007

21) A. Brandt, I. Kossev, A. Falke, J, Eickhoff, H.-P. Roeser, "Preliminary System Simulation Environment of the University Microsatellite Flying Laptop," Proceedings of the 6th IAA Symposium on Small Satellites for Earth Observation, Berlin, Germany, April 23 - 26, 2007

22) J. L. Jørgensen, P. S. Jørgensen, G. Grillmayer, "NEA detection, a possible use of the Flying Laptop microsatellite reconfigurability," Proceedings of the 5th IAA Symposium on Small Satellites for Earth Observation, April 4-8, 2005, Berlin, Germany

23) "The Flying Laptop Passed Its Review," IRS University of Stuttgart, March 9, 2017, URL:

24) "Flying Laptop — Academic Small Satellite Flying Laptop," IRS, 2017, URL:

25) Nico Bucher, Ulrich Mohr, Bastian Bätz, Kai-Sören Klemich, Sabine Klinkner, Jens Eickhoff, "Functional verification of the small satellite Flying Laptop," Proceedings of the 11th IAA Symposium on Small Satellites for Earth Observation, Berlin, Germany, April 24-28, 2017, paper: IAA-B11-0214

26) "ROSCOSMOS: Soyuz-2.1a Launch Vehicle with KANOPUS-V-IK Satellite Successfully Lifts Off From Baikonur," Roscosmos, July 14, 2017, URL:

27) Stephen Clark, "Soyuz rocket lifts off with 73 satellites," Spaceflight Now, July 14, 2017, URL:

28) "The small satellites 'TechnoSat' and 'Flying Laptop' are successfully launched into space," DLR, July 14, 2017, URL:


30) Julia Dancer, "The Flying Laptop is Being Launched from Baikonur on July 14, 2017!," IRS, University of Stuttgart, Germany, April 6, 2017, URL:

31) "Launch campaign for „Soyuz": maiden flight for our DCSM unit," Orbital Systems, May 25, 2017, URL:

32) Leena Pivovarova, "48 Doves to Launch on a Soyuz Rocket," July 7, 2017, URL:


34) "Russia will launch into orbit 72 small satellites per launch," Russian Aviation, 15 June 2017, URL:

35) M. Lachenmann, S. Walz, H.-P. Roeser, "Video Compression of the Flying Laptop for low Bandwidth Satellite Links," Proceedings of the 6th IAA Symposium on Small Satellites for Earth Observation, Berlin, Germany, April 23 - 26, 2007


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 (

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