The mechanical structure of the Flying Laptop is divided into the service module, the core module and the payload module as shown in the figure below.
The launch adapter is attached to the back plane of the service module. All modules are made of aluminium due to its high heat conduction properties and easy machining. In order to ensure the alignment of the visible and near infrared cameras (VIS/NIR) to each other and to the star cameras, all components are attached to an optical bench made of carbon-fiber-reinforced plastic (CFRP). The focus distance of the thermal infrared (TIR) camera and the Ka-band antenna is also influenced by thermal extension. Hence, the primary mirror and the retaining structure of the secondary mirror will also be produced from the temperature stable CFRP. For the TIR the primary mirror demands a medium surface roughness of approx. 0.8 µm.
The thermal system of the Flying Laptop is intended to be passive by using the dissipated heat of the internal components. The surface is mainly covered by multi layer insulation (MLI) and the heat is released by a radiator on the back plane of the payload module.

Modular design of the Flying Laptop, a) Payload Module, b) Core Module,
c) Service Module

Power is generated by three solar panels with triple-junction solar cells. Two panels are deployable. The battery pack consists of 8 NiH2 cells and can deliver the required peak power consumption of 300 W for the travelling wave tube.

The satellite is 3-axis stabilized by actuators, which consist of reaction wheels and magnetic torquers. The four reaction wheels are aligned in a tetrahedron configuration.and the three magnetic torquers (torque rods) dump the momentum accumulated by the reaction wheels. The moment of inertia in the x, y and z axis of the satellite is estimated to be around 4 kgm².
The attitude motion is monitored by five different types of sensors: a 3-axis magnetometer, a 4-pi coarse sun sensors system, rate sensors, a star tracker unit with two camera heads and GPS receivers. The Zarm AMR-magnetometer uses a magneto-resistive sensor and has a digital interface. For the measurement of the angular velocity, four fiber optical rate sensors will be used. A star tracker, the micro Advanced Stellar Compass (µASC), from the Technical University of Denmark will provide a relative pointing knowledge of up to 2 arcseconds. When the satellite is stabilized and rotates with a slew rate of less than 1.2 °/s the star tracker delivers regular attitude updates. To provide full accuracy about all axes and to decrease the probability of blinding during maneuvers, a second camera head unit is mounted on the satellite with its optical axis tilted away from the first one. To support accurate target-pointing of the spacecraft during imaging and ground station contacts, the satellite will be equipped with a GPS navigation system consisting of three Phoenix receivers from DLR/GSOC.

Imaging Modes: A) Inertial- , B) Nadir- & C) Target-Pointing Mode

For image acquisition three different attitude control modes are defined and shown above: inertial-pointing mode, nadir-pointing mode and target-pointing mode. In the target-pointing, also known as spotlight mode, the satellite points to a fixed spot on the surface of the earth during a fly-over. This allows longer integration times for the cameras which is a significant advantage for the scientific measurements. The slew rate for this maneuver is 1 °/s (max.) and follows a non-linear bell-shaped curve over time. This is the most demanding mode of the satellite in terms of control algorithms.

Beside the systems for technology demonstration the Flying Laptop is equipped with two imaging payload instruments. For the measurement of the BRDF a multispectral camera system is implemented. This system consists of three single cameras, one for each channel (green, red and near infrared), mounted in a triangle mode. The system is directly controlled from the OBC which also serves as data handling unit and will take images with a ground sample distance (GSD) of 25 m (1024 x 1024 Pixel). The resolution was chosen for co registration of the thermal infrared images. All cameras are mounted on a reinforced carbon fibre composite plate for better alignment and thermal stability. Also mounted on this plate are the star cameras for precise orientation of the imaging system.
The second payload is a thermal infrared camera system for images with a GSD of 50 m (320 x 240 Pixel). The camera consists of a cassegrain mirror system (dual use system described in the section technology demonstration), a relay optic and a micro-bolometer sensor. Therefore no cooling is necessary which makes the system small in size with low power consumption and also cost effective.

The on-board computer (OBC) consists of a 4-node FPGA computer.

For telemetry and telecommand VHF, UHF (low gain) and S-band (low and high gain) antennas will be installed on the satellite. Beside S-band communication, VHF and UHF offers the possibility to utilize amateur radio equipment. The 50 cm mirror of the Flying Laptop works as a dual system. On one hand, it is the antenna reflector for the Ka-band communication; on the other hand, it is the primary mirror for the optical system of the thermal infrared camera. The two wavelengths are separated with the help of a beam splitter. By combining these two complex systems a compact design is achieved which can be flown on a micro-satellite. The use of a TWT is an unique feature of this micro-satellite.

Setting up a verification environment for reliable system-wide tests is new to micro-satellite projects, but it is one of the enabling technologies for proving the required attitude control system accuracy. In this context a software-based functional verification reduces the check-out environment complexity and huge costs can be saved.
This model-based verification environment for small satellite applications is under development in close cooperation with EADS Astrium and will be set up in parallel to the Flying Laptop development. It is characterized by high real-time capabilities to represent the spacecraft hardware system-wide in its exact operational modes and response times. Software models of the spacecraft components will be created successively in adequate detail in order to provide the particular test bench functionality. Latest commercial improvements in hardware and software technology allow the real-time test benches to be set up using standard computers and a Linux operating system kernel which supports real-time performance.
A Software Verification Facility is in progress to support the on-board software development process. It is used to debug, validate and verify the on-board software and prove the overall system dataflow functionality. Consisting of a real-time simulator, an on-board computer simulator and a central control system, it supports real-time or accelerated software simulation of the whole spacecraft system by implementing all components by its hardware-specific software models into the simulation environment.
In the next step a FlatSat test bench with real-time performance will be arranged as first hardware check-out environment including only the on-board computer as hardware in the loop. It will be set up well before spacecraft integration using a test harness to maintain single component up to system-wide check-out procedures. All component software models, especially those of the attitude control system, will be re-used in this environment to verify correct operation.
The third test bench is created as an expansion of the FlatSat test bench and integrates additional spacecraft hardware in the simulation loop. Single component check-out tests will be followed by complete mission scenario simulations. Finally the protoflight test bench supports a functional verification test environment throughout the flight hardware qualification process.

RTB Test Bench (source: EADS Astrium)