Command & Data Handling (CDH)
The command and data handling (CDH) system design is driven by the necessity of low power usage, reliability, flexibility, simplicity of development, and recognition that a Cubesat is essentially an autonomous agent with occasional periods of supervision. It must be capable of taking care of itself and possess a limited capability to deal with unplanned contingencies or at least enter a fail-safe mode leading to a recoverable state for which communication can be used to diagnose problems from the ground. Our CDH is driven by a TI MSP430 microcontroller unit on Pumpkin Inc's FM430 board.
TopCommunications System (COMMS)
The success of the mission relies heavily on the ability to transfer information between the satellite and the ground station (i.e. mission control). To expedite the development process and establish a realistic launch schedule, a commercial-off-the-shelf radio board from The Stensat Group LLC will be used for this mission (see Figure 5-4). The Stensat radio board provides a fully integrated communications transceiver based on the AX.25 packet data standard. Radio communications will take place over amateur radio channels within the VHF (uplink) and UHF (downlink) frequency bands using narrow band frequency modulation (FM).
Subsystem Architecture
Transmit (downlink) data is sent to the Stensat radio board via I2C from the MSP430 processor in the Command and Data Handling subsystem. The signal is then modulated up to RF by means of the Stensat TX-51 transmitter module, amplified, and filtered, all onboard the Stensat radio board. Next, the signal will go through an antenna matching unit for impendence matching (if necessary) before being sent to the transmit antenna, mounted externally. Received (uplink) data follows a similar but reversed path. The RF signal is obtained at the receive antenna, which connects to an antenna matching unit for impendence matching (if necessary). The signal then passes through a filter and preamplifier before going to the MC13135DW FM Receiver, all onboard the Stensat radio board. The received data is then sent to Command and Data Handling across the I2C bus. The current used at the transmitter and receiver is monitored by two ZCXT1022 current sensors (i.e. one each for the transmitter and receiver). Each sensor then sends analog current data to an ADF7012 Analog-to-Digital converter (i.e. 2 A/D converters total). Digital current data is then routed to the Health Management system.
Uplink/Downlink Specifications
As previously mentioned, the both uplink and downlink use the AX.25 protocol for packet data transmission. The AX.25 protocol allows up to 200 characters per packet. Only text, punctuation, and numerical ASCII characters are allowed. The uplink (ground to satellite) will operate in the VHF band at a carrier frequency of 145.980 MHz. The audio frequency-shift keying (AFSK) modulation scheme is used, allowing for a receive data rate of 1200 baud (i.e. 1200 symbols/second). Received signal sensitivity requires a total received power level of -105 to -110 dBm. The downlink (satellite to ground) will operate in the UHF at a carrier frequency of 437.385 MHz. The transmitter can operate in two modes: 1200 baud AFSK and 9600 baud FSK (frequency-shift keying). The 1200 baud AFSK mode will be used as a beacon to establish ground communication. Once the link is established, the transmitter will switch to 9600 baud FSK mode for increased data transfer rates on the downlink. The transmitter consumes 1 W of power while active, which is included in the power budget. The antenna design for both transmit and receive antennas is yet to be determined. However, common designs for cubesat antennas include dual half-wavelength dipole antenna or dual quarter-wavelength monopole antenna arrays. Each typically utilizes one antenna for uplink and one antenna for downlink. The final design must meet total radiated power (TRP) and total isotropic sensitivity (TIS) requirements based on the AX.25 protocol to ensure proper data communications. Additional research into antenna design is required. The frequencies used for telemetry were granted by the International Amateur Radio Union on January 14, 2007.
Mission Control
The communications chain is completed by the ground control station. This need will be filled by UF's Gator Nation Earth Station (GNES) ground station, which is equipped to copy and receive AX.25 digital packets. The GNES User's Manual has been updated and is available for download at the Gator Amateur Radio Club's website. If possible, the equipment and expertise of NCSU's W4ATC Student Amateur Radio Society will also be utilized to provide a larger footprint (communications window), as well as redundancy if one of the mission control systems becomes non-functional. The term "communications window" refers to the period of time during which a direct line-of-sight communication channel is available between the satellite and the ground station. The exact length of the communication window will be dictated by the specific orbital geometry of the satellite, which is the result of launch parameters defined by the launch provider. Based on initial calculations, the communications window between a Low-Earth Orbit (LEO) satellite and a single ground station should last approximately 5- 10 minutes and occur approximately 4-6 times per day. Thus, the total daily communications window is approximately 20 to 60 minutes. The success of the mission hinges heavily on the communication system. Instructions from the main MCU will be sent to a digital transceiver via the I2C bus. The transceiver uses the AX.25 protocol for packet transmission and is capable of producing downlink transmissions at 1200 baud AFSK and 9600 baud FSK on 437.385 MHz. Uplink transmissions are received using 1200 baud AFSK on 145.980 MHz. The communications chain is completed with the Gator Nation Earth Station (GNES) ground station which is equipped to transmit and receive digital packets. The link budget used to complete the radio link is available upon request.
Attitude Control System (ACS)
Attitude Determination System (ADS)
In order to effect attitude control it is necessary to be able of determine the attitude of the satellite. The SGCMG payload has a theoretical pointing precision an order of magnitude greater than the expected accuracy of our attitude determination. While real-time attitude determination using a star tracker is not feasible for this mission, the use of a camera for post processing attitude determination is being studied. An image of the sky before and after an attitude maneuver will be taken and stored. The images will be transmitted to the ground and attitude for each position will be determined on the ground. Knowing the desired change in attitude and the actual change in attitude will provide a more precise measure of the precision achieved by the maneuver.
Attitude will be estimated from measurements of the sun vector and the Earth's magnetic field vector using sun sensors and a magnetometer. Models of the magnetic field and sun vector stored on board will be used to relate the body frame of the satellite to the Earth Centered Inertial frame and estimate the attitude. A GPS receiver will be incorporated to provide the satellite's position for the calculations.
Electronic Power System (EPS)
The power subsystem is responsible for gathering, storing and distributing power to the
cubesat system. Power gathered by the solar cells will be stored in Li-Po battery packs
and distributed through the electrical power system (EPS). The EPS will host a
dedicated microcontroller to monitor the various circuitries and energy sources. Solar
array voltages and temperatures, MPPT currents, and battery capacitance will be
monitored by the MCU on the EPS and the information will be communicated to the
CDH module through the I2C interface.
Power Generation & Management
The block diagram depicts the layout of the EPS to be used to
distribute power to the cubesat system. Some of the components and circuits included
in the EPS are discussed below.
Triple Junction Solar Cells
Triple junction photovoltaic cells will be used to harness energy from the sun and use it to power the cubesat system. Solar cells from Spectrolab are capable of supplying voltages up to 2.2V per cell and two such cells connected in series will be mounted on each face of the cubesat system to produce 4.4V. The solar arrays on the 6 faces of the cubesat will be connected in parallel and the power generated will either be stored in the Li-Po battery packs or will be supplied to the power bus through the voltage regulators. The solar arrays will be monitored by the Microcontroller Unit (MCU) of the EPS through hall type current sensors. These sensors also provide sun vector information for the attitude determination system.
Maximum Power Point Tracking
In an architecture where the battery packs are being directly charged from an array of solar cells the performance of the solar arrays is at its maximum when Š i) the battery packs are fully charged and ii) the temperature of the solar panels is at a maximum. When the cubesat enters the sunlit path after an eclipse, the solar array temperature is low and the array voltage is clamped to the battery voltage. During this time the energy extraction from the solar panels is highly inefficient. Similarly, the energy extraction from the solar arrays is maximum when the cubesat is at the end of the sunlit portion of the orbit. To optimize and ensure maximum power generation during the entire orbital cycle MPPT circuits will be used on each solar array. In the absence of MPPT circuits solar cells tend to produce power at a voltage which is not ideal for efficient power production. To limit the voltage and extract maximum power from the solar cells each array will be guarded by an MPPT circuit. The MPPT circuits will be powered by the solar arrays.
Polymer Lithium Ion Battery Packs
Power produced from the solar cells will be stored in Li-Po battery packs which will power the cubesat system when the solar cells are not generating power. Li-Po battery packs have been chosen for their high energy density and their past use in space applications. Two Li-Po 3.7V cells connected in series will make up each battery pack. Fuel gauges connected to each battery pack will monitor the capacity and communicate it to the MCU. The typical Orbital Average Power (OAP) of a 1U cubesat in a sun synchronous orbit is around 2W. Assuming the Depth of Discharge (DoD) of the batteries to be 20% the Li-Po batteries considered here will power the cubesat for six hours.
Charging Circuit
Power from the solar panels will be routed to the Li-Po battery pack through a charging circuit. To protect the batteries from over-voltage under-voltage, overcharge current and over discharge current the charging circuit will also include a Li-Po battery protector. The battery protector will disable the output of the 5V and 3.3V converters once the battery reaches its minimum acceptable level. The output of the converters will be enabled once the batteries reach an optimum acceptable level.
Voltage Regulators and Switching Circuit
Since all of the subsystems require either +5V or +3.3V supply, the power from the solar panels and the Li-Po battery pack will be routed through voltage regulators or DC/DC converters. An intelligent switching circuit will be used to switch between solar panels and the battery pack. The EPS MCU will have access to the switching circuit and will be capable of overriding the circuit if required.
Power Budget and Balance
Estimating the amount of incoming power and outgoing power is important while planning the mission pay load. The incoming power is a function of attitude of the cubesat. The power generated in the solar cell depends on the angle of incidence of the solar rays or the cubesatÕs angle with sun. The cubesat is illuminated on either zero, one, two or three sides simultaneously. The cubesat will orbit the earth 16 times a day with an orbital period of 90 minutes. Each orbit has a sun time of 60 minutes and has an eclipse time of 30 minutes. The average load power available from the sun is a useful number to have.
Health Management System (HMS)
In order to mitigate component failure or unforeseen changes due to the space environment, the satellite's operating system will periodically poll certain components to determine their level of functionality and the overall health of the spacecraft. Specific quantities to be monitored include
System health monitoring sensors such as hall-effect current sensors, temperature sensors and watchdog timer will communicate with the CDH module directly or through the microcontroller unit on the EPS to ensure proper functioning of the cubesat system. Hall type current sensors will be used to monitor the functioning of the flywheel and gimbal motors. Batteries will be monitored by the EPS MCU using fuel gauges. A watchdog timer will be used to monitor the functioning of the onboard processing units. Temperature sensors will be mounted on different faces of the cubesat structure. The sensors will be interfaced to the processing unit through A/D converters.
Thermal Considerations
The thermal analysis ensures that the operation of the satellite can continue during extreme conditions by utilizing a combination of heaters, temperature sensors and insulation techniques. Every component will be chosen to meet the expected temperature range of -40 to 50 degrees Celsius. In the event that external heating will be needed for operation, the health monitoring system will trigger the heaters for optimal involvement. Temperature sensors will be intermittent within the critical items such as batteries and circuitry temperatures. Where necessary, insulation will be implemented. One method of insulation is to apply temperature resistant resins to component surfaces. The components will undergo tests in a thermal vacuum chamber. This testing will provide a rigorous approach to the overall selection of hardware and health monitoring sensors. ANSYS is a finite element program that will be used to evaluate thermal and structural models of the satellite.