| Auto: | all |
| SumbandilaSat | |
| Mission Type: | Technology |
| Operator: | University of Stellenbosch |
| Cospar Id: | 2009-049F |
| Satcat: | 35870 |
| Mission Duration: | 3 years (planned); (achieved) |
| Manufacturer: | SunSpace |
| Launch Mass: | 81kg (179lb) |
| Launch Date: | UTC |
| Launch Rocket: | Soyuz-2-1b/Fregat |
| Launch Site: | Baikonur Site 31/6 |
| Disposal Type: | Damaged by solar storm |
| Decay Date: | 10 December 2021[1] |
| Orbit Epoch: | 24 May 2015[2] |
| Orbit Reference: | Geocentric |
| Orbit Regime: | Low Earth |
| Orbit Periapsis: | 447.7km (278.2miles) |
| Orbit Apoapsis: | 449.8km (279.5miles) |
| Orbit Inclination: | 97.1 degrees |
| Orbit Period: | 93.4 minutes |
| Apsis: | gee |
SumbandilaSat (sometimes just Sumbandila, formerly ZASAT-002, AMSAT designation SO-67[3] [4]), was a South African micro Earth observation satellite, launched on 17 September 2009 on a Soyuz-2 launch vehicle from the Baikonur Cosmodrome.[5] The first part of the name, Sumbandila, is from the Venda language and means "lead the way".
The University of Stellenbosch, SunSpace and the CSIR (Council for Scientific and Industrial Research) were key players in constructing SumbandilaSat. The CSIR's Satellite Application Centre (CSIR-SAC) was responsible for operations, telemetry, tracking, control as well as data capturing.
SumbandilaSat was part of a closely integrated South African space programme and served as a research tool to investigate the viability of affordable space technology. Furthermore, the data was used to, amongst others, monitor and manage disasters such as flooding, oil spills and fires within Southern Africa.
In June 2011 the satellite was damaged during a solar storm. The damage caused the on-board computer and the camera to stop functioning. This has caused it to stop fulfilling its primary objective and it has been written off as a loss by SunSpace, its builder.[6] The final data packet from the satellite was received on 14 September 2011.[7]
The launch site was the Baikonur Cosmodrome LC-31/6, which is located at the following coordinates:
| Imager | NER < 0.6% with a forward motion compensation (FMC) factor of 4:1 |
|---|---|
| Operational MTF: >= 5% over the full field (excluding orbit motion effects) | |
| GSD = 6.25m @ an orbit altitude of 500 km | |
| 6 spectral band (visible range) line scanner | |
| Matrix sensor included for "snapshot" pictures | |
| Image quantisation: 12-bit (data for each pixel stored as 2 bytes) | |
| Image modes | Default scan mode with FMC = 4:1 but system could operate with FMC = 1:1 with consequent degradation in NER |
| In FMC = 4:1 mode, non-contiguous scenes of 45 km x 45 km could be imaged (max 10 scenes in 6 spectral bands before data downloading was required) | |
| In FMC = 1:1 mode, a contiguous strip with 45 km swath could be imaged (max track length of 450 km could be imaged in 6 spectral bands before data downloading was required) | |
| Imager data store | 24 Gbyte |
| Image data downlink | Expected frequency implemented on satellite exploration S-band |
| Link margin: 3 dB @ 10° (calculated with 0dBi satellite antenna; 5W Tx power and SAC GS parameters) | |
| Data rate sufficient to download full image data store during two night passes | |
| No real-time downloading of images (all images stored on board the satellite) | |
| Viewfinder | Live downlinking of PAL video images during TT&C ground station passes |
| PAL images selectable between B&W (narrow FOV) and two wider FOV colour PAL cameras | |
| Satellite bore-sight steerable with "joystick" interface | |
| Viewfinder mode could be interrupted with either image snapshot or image linescan modes upon ground command | |
| The bore-sight direction of the viewfinder was the same as the main imager | |
| ADCS system | The satellite was 3-axis stabilised |
| System performance was sufficient to maintain pointing accuracy for image downloading | |
| ADCS performance did not degrade image quality | |
| The satellite bore-sight could be controlled to within 3 km on the ground | |
| The intended system implementation used a combination of the following actuators and sensors: Horizon, fine-Sun, coarse-Sun sensors; star camera; magnetometer(s); fibre-optic gyros; reaction wheels and magneto torquers | |
| A satellite slew manoeuvre from one stabilised position to another stabilised position, through an angle of 30°, could be completed in less than 1 minute | |
| Propulsion System | Sufficient propellant included to maintain a satisfactory orbit for 3–4 years |
| System was to demonstrate orbit constellation deployment as well as orbit maintenance | |
| TT&C Communication Link | Expected frequencies implemented on commercial VHF uplink and UHF downlink |
| Link margin: 6 dB @ 5° (calculated with -12dBi satellite antenna null; 5W Tx power and 12dBi GS antenna gain) | |
| Satellite Housekeeping | One TT&C GS pass per 24h was sufficient for the purpose of monitoring telemetry and uploading of new command sets and SW |
| Power | Energy source: solar panel with 65W (EOL) capacity |
| Mission planning dictated energy requirements per orbit | |
| Experiments | Provision was made for two 1-kg experiments |
| SU certified the space environmental readiness of the experiments at MC level prior to integration | |
| Average power available per experiment: 1.2W (TBC) | |
| Peak power available per experiment: 10W (TBC) | |
| Orbit lifetime | Design lifetime of 3 years at an orbit altitude of 500 km (subject to average Sun activity) |
| Given the ultimate unpredictability of the space environment, the operational life varied from the design lifetime |
It had a number of secondary experimental payloads on board: