SailSat-1

1. A 3U CubeSat for the QB50 Mission SailSat-1 Team Members: Matthew HorganAjith SasidharanMichael RodwellElliot WertheimerTrent Jansen- Sturgeon

2. Admire the artistry...

3. AGAIN...

4. The Payload

5. Materials * Both materials are coated with Aluminium, and hence have the same reflectivity. ** By the test method of “ASTM E-794-85 (1989)”.

6. Folding Techniques Figure 1: The "Leaf-out" pattern as described by de Focatiis and Guest (2002) (Surry, 2013).

7. Deployment Figure 2: "Tape-measure" supports (Adeli, 2012).

8. On Board Data Handling Subsystem

9. Technical specifications • ARM7 processor - 8-40 Mhz • 2MB RAM, 4MB Code storage, 4MB data flash • CAN and I2C interfaces • MicroSD socket for up to 2GB storage • 3x PWM drivers • 6x sun-sensor inputs • 3x Rate-gyro inputs Unchanged from PDR

10. OBDH • Layered format • Constant communication with other subsystems

11. OBDH Interfaces • Primary interfaces are CAN and I2C • CAN is real time, secure, faster • I2C Supports bi-directional data transfer

12. ADCS-OBDH • Connects to magnetorquers and reaction wheels using 3-PWM bidirectional outputs • Analog interface to connect to photo diodes

13. OBDH schematic

14. Physical properties • 96mm x 90 mm x (16-26)mm • Mass: 50-55 grams • Operating temperature range is -40 to 60 °C

15. Power Subsystem

16. Power Subsystem • Combination of Solar Panels & Batteries • All components from ClydeSpace

17. Solar Panels • The solar panels we will be using are the ClydeSpace 3U CubeSatsolar panels. • Four ‘Side Solar panels’ (SP-L-S3U-0016-CS) and only one ‘Top/Bottom Solar panels’ (SP-L-T1U-0017-CS) because one face will have the deployed solar sail attached via a boom.

18. Batteries • ClydeSpace Standalone Lithium Polymer battery or CS-SBAT2-30 • The EPS is fully integrated with 3 lithium polymer batteries in parallel, providing approximately 30Whrs of capacity. • The choice of lithium polymer batteries and Clyde-Space products was for the maximum utilization of internal spacecraft volume

19. EPS • 3U ClydeSpace EPS • Power from the solar array sections are transferred to the battery bus via battery charge regulators

20. Power System Schematic

21. Power Balance

22. Communication Subsystem

23. Link Budget • Equivalent Isotropically Radiated Power (EIRP) • Free space path loss (FSPL) • Atmospheric Loss • Rain loss • Antenna point loss • Ionospheric loss • Transmission line loss

24. DOWNLINK

25. UPLINK

26. FIGURE OF MERIT • The ratio of receiving antenna gain to thermal noise: • (dB) • Uplink thermal noise approx. 350K • Downlink thermal noise approx. 610K • Difference due to very low CubeSat antenna temperature (30K)

27. Carrier to Thermal Noise Ratio • All gain and loss of signal strength considered together:

28. Carrier to Noise Density • Gives an indication of the radio of strength of signal to thermal noise density • It is found by diving the C/T ratio by the Boltzmann constant or in log terms: The Boltzmann constant is 228.6 dBW/K*Hz

29. Energy per Bit to Noise Density • This is the normalised ‘signal to noise ratio’ • The final calculation in determining whether uplink/downlink communication is viable with our design ‘R’ is the bit rate (1200kbps for uplink and 9600kbps for downlink)

30. LINK MARGIN • According to provided data, and considering ground station and CubeSat modulation procedures: Downlink margin  13.55 dB Uplink margin  30.27 dB *These values are well within safe design boundaries and have also accommodated for the worst possible and unlikely scenario.

31. PCB Design

32. PCB Design

33. Antenna

34. Attitude Determination & Control Subsystem

35. Sensing Systems • 6 sun sensors (Cubesatshop.com): • Field of view: 114° • Power: < 0.05W • Mass: <5g • Size: 33mm x 11mm x 6mm • Accelerometer (ADXL335): • Operating current of 6.5 mA • Supply voltage of 2.1 to 3.6V • Mass: 15g • Dimensions of 17.78mm x 13.97mm • Gyroscope (ITG-3200): • Sensing range +/- 3g • Supply current of 320 μA • Supply voltage of 1.8 to 3.6VDC • Mass: 10g • Dimensions of 17.78mm x 17.78mm Obtain the sun vector and solar influx. Linear acceleration of the spacecraft in three axes. Rotation rate of the spacecraft in three axes.

36. Actuating Systems • 3 axis reaction wheel box (MAI 101): • Maximum torque: 0.635mNm • Momentum storage: 1.1 mNms • Power supply: 12 to 28 VDC • at 200 mA(360 mAmax) • Mass: 640g • Dimensions: 76.2mm x 76.2mm x 69.85mm • 3 Magnetorquerrods (Cubesat shop): • Magnetic moment: 0.2 Am2 • Power supply: 200 mW at 5V • Diameter: <9mm • Length: 7cm • Mass: 30g Modify the spacecraft orientation. Desaturate the reaction wheels or detumble the spacecraft.

37. Components Interaction Communications and data storage Magnetorquers Sun sensors Onboard computer and software treatment Linear accelerometer Reaction wheels Rate gyroscope

38. Controllability α=T/I T=Iα Folded antenna and sail configuration: Fully deployed configuration:

39. Mass Budget

40. Thermal Subsystem

41. Thermal Requirements • Battery: -10 to 20̊ C • Power Regulating Unit: 0 to 40̊C • Electronic Power System: -40 to 85 ̊C • Reaction Wheels: -5 to 45̊C • Attitude Determination Electronics: 0 to 40̊C • Antenna Elements: -100 to 150 ̊C • Inactive Structures: -100 to 100 ̊C

42. Thermal Subsystem • Passive Control • Saves Space, Mass, Power • MultiLayer Insulation (MLI) Blankets • Polyimide Film developed by DuPont [Kapton] • Reduce temperature swings endured by components caused by orbit transitions from the sunny to the shaded areas. • Durable

43. I hope you enjoyed our presentation and if you did not, I hope you had a good nap.