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DubaiSat-1 Technical Specifications

DubaiSat-1 Satellite Technical Specifications

The DubaiSat-1 Satellite is designed to be compatible with low earth sun synchronous orbit at an altitude of 680km

Technical Specifications

  • The DubaiSat-1 Satellite is designed to be compatible with low earth sun synchronous orbit at an altitude of 680km.
  • DubaiSat-1 with its stowed solar panels is compatible with most small satellite launchers as a secondary payload or as a primary payload such as Dnepr.
  • The spacecraft bus provides a typical pointing performance of 0.2 deg.
  • DubaiSat-1 weighs less than 200 kg including 50 kg payload mass.
  • The average power consumption is less than 150 Watts.
  • DubaiSat-1 features a deck-and-longeron type structure, allowing easy assembly and disassembly with about 1.2 metres in diameter and 1.35 metres in height.
  • Most of the component units are positioned on the bottom deck, while reaction wheels and gyros are positioned on the middle deck.
  • The interface with the launch vehicle is made through an adapter (specific to each launch vehicle) bolted to the bottom of the structure. The mechanical interface with the payload Electro-Optical Subsystem (EOS) is provided through three points at the middle deck.
  • The satellite was designed for a minimum lifetime of 5 years in-orbit operation.

 DubaiSat-1 Satellite consists of the following parts:

1. Command & Data Handling (C&DH)
The Command and Data-handling Subsystem is used to control the flow of communication between all modules in the satellite, including the payload. It uses two On-Board Computers (OBC) - one as a redundant, four Telemetry and Command Modules (TCM) - two as redundant, a GPS module and the On-Board Software.

The OBC uses ERC-32 processor, which is used in many ESA missions. It includes the FOS which controls the telemetry and commands, saves them and performs scheduled tasks with different satellite subsystems. It also includes the Flight Control Software (FCS) which is specified for controlling the satellite attitude and performs maneuvers if necessary.

Two TCMs are used to collect data from subsystems and control the flow in and out from the OBC. The GPS module is used for time synchronization between modules and for fine satellite positioning.


1 Figure ‎(1) shows the functional Block diagram of C&DH subsystem with all other subsystems connected to it:


2. Attitude Control System (ACS)
The Attitude Control Subsystem is used to both monitor and control the satellite attitude during the full mission lifetime, except during launch. It is capable of tilting the satellite to 45º along and across track. It has a pointing accuracy of < 10 arc sec and control stability of better than 0.016 º/sec. The commands are generated in the OBC using the Flight Control Software (FCS).

The system consists of the following sensors; Fine and Coarse Sun Sensors (FSS & CSS), Magnetometers (MAG), Fiber-optical Gyros (Gyro), and Star Sensors (STS). As actuators, it has four Reaction Wheels (RW) and Magentorquers (MT). The four RWs are configured in a tetrahedral scheme to provide optimum performance redundancy. They will be desaturated by MT commands.

The ACS components are all space qualified and have been used on previous missions successfully.

3. Electrical Power System (EPS)
The Electrical Power Subsystem in the satellite generates, stores, regulates, and distributes electrical power to all subsystems including the payload. It consists of three Solar Arrays, three Battery packs, and Power Electronics.

The solar arrays generate power for all the satellite subsystems and batteries, and are attached on three deployable solar panels. NiCd batteries are used as energy storage devices for eclipse operation and to compensate for excessive power consumption in high power consuming operations during daytime.
In normal operation, the power from the solar arrays is used to provide > 330W (EOL) in sunlight. This power is distributed to all subsystems in the satellite and also used to charge the batteries. During eclipse period, power from the batteries is used to all subsystems with maintaining low performance of subsystems. The Depth of Discharge (DoD) of all three batteries < 20%.

The solar arrays and batteries are space qualified and have been used on other missions successfully.

The BOL and EOL power values depend on temperature of operation. For DubaiSat-1, the solar panels will be operating at an average of 65 degrees Celsius. At this temperature:

Power generated at BOL (3 solar panels) = 360W
Power generated at EOL (3 solar panels) = 336W

For DubaiSat-1, the nominal battery capacity at BOL (for 3 batteries) is stated as 18Ah (actually around 22Ah when tested). The EOL battery capacity is considered 14.4Ah.

4. Structure & Thermal Mechanical:

4.1 Structure Mechanical: 
DubaiSat-1 is a hexagonal shaped satellite with three deployable solar panels. The coordinate system is defined as follows: +X is defined along the scanning direction of DMAC, which is normal to the detector line, +Z is aligned with the DMAC optical axis, and +Y is the deployed direction of one of the solar panels (SP-C) and perpendicular to +X and +Z axes. The origin of the coordinate is defined at the centre of the adaptor.

Three S/C adaptors are connected to the shear brackets and mechanically couple the satellite and separation adaptor of the launch vehicle. Each one of the S/C adaptors has two S/C-based separation sensors and one LV-based separation sensor to monitor separation condition between the satellite and the launch vehicle.


Figure (2) shows the configuration of the DubaiSat-1 satellite when the solar panels are folded


Figure ‎(3) shows the configuration of DubaiSat-1 satellite when the solar panels are deployed.


4.2 Thermal:
DubaiSat-1 bus system utilizes passive thermal control, while DMAC payload and battery system implement active thermal control methods. All bus subsystems are thermally interconnected. DMAC is thermally isolated from the bus.

Operational heater for EOS is controlled by TPM and has full redundancy. Pyro & Heater Control Electronics (PHCE) controls other heaters. Especially, Survival heater for battery is operated during the sunlight period only.

5. Telecommunication System (TS)

The satellite has two types of telecommunication systems for the contact with the Satellite Imaging. S-band transmitter and receiver are used for telemetry and command. X-band transmitter is used for image data transmission to the ground.

Two S-band transmitters (STX) are used for redundancy. The STX has output power of 2 W, which secures +33 dBm for data link. Two S-band receivers (SRX) are also used for redundancy. An MMIC is used as a low-noise amplifier LNA to amplify weak received signal as it has low noise and high gain characteristics with high reliability. Two S-band Antennas are placed on the top and bottom of the satellite to provide omni-directional coverage. The SRX and STX share these antennas using a duplexer and a power divider.

Two Image Transmission Units (ITU) are used for redundancy. Each produces a 5 W (+37dBm) signal for high data rate transmission using QPSK modulation.

6. DubaiSat-1 Medium Aperture Camera (DMAC)

The payload onboard DubaiSat-1 is an optical camera used for imaging. It is a push-broom imaging system with one panchromatic and four multi-spectral imaging channels in the Red, Green, Blue and NIR channels in the following wavelengths, PAN 420nm – 720 nm, Blue 420 – 510 nm, Green 510 – 580 nm, red 600 – 720 nm and NIR 760 – 890 nm. The panchromatic resolution is 2.5m and the multi-spectral is 5m resolution. DMAC consists of the Electro-Optical Subsystem (EOS) and the Payload Management Subsystem (PMS).

EOS consists of the Telescope, the Focal Plane Assembly (FPA), and the Signal Processing Module (SPM). PMS consists of the Thermal & Power Module (TPM) and the Mass storage & Control Module (MCM).

The Telescope is made of various optical and mechanical components. It is an on-axis catadioptric telescope shared between all imaging channels of the DMAC system. It collects the incoming radiation from ground targets and focuses the collected radiation onto FPA. It consists of Primary Mirror Assembly (M1A), Secondary Mirror Assembly (M2A), Correction Lens Assembly (CLA), Metering Structure, and other structural components.

The FPA consists of linear detectors, which convert the focused radiation into raw video signal in five imaging channels. The SPM receives the video signal and processes it and convert it into a digital signal sent to the MCM.

The MCM module receives the digital signal and stores it for direct download or scheduled download. It also manages the temperature control of the payload and the power supplied to the electronics and sensors.

All the modules have redundancy except the optical part and the FPA.


Figure ‎(4) shows the schematic of DMAC internal configuration as a functional diagram:


7. Space Radiation Monitor (SRM)

The Space Radiation Monitor is capable of measuring the total ionizing dose from the population of charged particles at the orbits of satellites. The instrument utilizes four (4) p-type Metal-Oxide-Semiconductor Field Effects Transistors (MOSFETs) devices to measure the current-voltage (I-V) characteristics. Subsequent analysis of the I-V will allow measurements of amount of cumulative ionizing dose, Total Ionizing Dose (TID), of the devices. A simplified block diagram is shown in Figure ‎(5).