UniSat (University Satellite)
UniSat (University Satellite) Program
UniSat is an experimental microsatellite program of the University of Rome (Universita di Roma “La Sapienza”, Scuola di Ingegneria Aerospaziale). Education and hands-on experience of students and faculty in the field of aerospace engineering is the overall theme of the program (involving a number of Italian universities), funded by the Italian Government MURST (Ministry of University and Technological Research) and by ASI (Italian Space Agency). In the framework of the UniSat program, a laboratory for satellite design and construction and an amateur ground station called SPIV (San Pietro in Vincoli), were established at the University of Rome. 1) 2) 3)
A launch of UniSat-1 took place on Sept. 26, 2000 from the Baikonur Cosmodrome on a Dnepr launch vehicle of ISC Kosmotras. UniSat-1, along with MegSat of Italy and TiungSat 1 of Malaysia, and SaudiSat-1A and -1B of SISR (Saudi Institute for Space Research), was launched in a multiple launch configuration (no primary payload). UniSat-1 is being operated by students.
Orbit: Circular orbit, altitude = 650 km, inclination = 65º, period = 98 minutes.
UniSat-1 is no more operational.
The power subsystem represents two experiments; one involves a solar panel environment test, the second experiment involves a functional NiMH (Nickel Metal Hydride) battery test. The new batteries offer a much better capacity/mass ratio than conventional NiCd batteries. The OBC distributed network architecture is a proof-of-concept test.
UniSat-2 (University Satellite-2)
UniSat-2 is a follow-up microsatellite project of the University of Rome (Universita di Roma “La Sapienza”, Scuola di Ingegneria Aerospaziale). The overall objective of this mission is to develop, and demonstrate an onboard micro-propulsion system, followed by a performance assessment. Such a system might eventually be used for future precise attitude control of a nanosatellite. 5)
UniSat-2, built by the GAUSS Laboratory of the University of Rome, is a spin stabilized spacecraft with the same dimensions as those of UniSat-1 consisting of an octagonal prism of 150 mm side length and 250 mm in height. The S/C has a modular structure made of five Al/Al honeycomb plates, kept together by four aluminum bars. Eight Al/Al honeycomb lateral panels, with surface-mounted solar panels, complete the satellite cylinder structure.
The S/C architecture consists of a multipurpose bus, able to accommodate several kinds of payloads (a tray system), and to accommodate change. The overall design permits easy assembling and disassembling, where each plate is devoted to a subsystem, thus allowing separate testing just before final integration. A Li-ion battery is used for ecliptic-phase operations, the S/C mass is 12 kg.
A launch of UniSat-2 from the Baikonur Cosmodrome on a Dnepr-1 vehicle took place on Dec. 20, 2002.
The co-passenger payloads on this multiple S/C flight were:
• LatinSat-1/2 (each of 12 kg) communications satellites for Aprize Satellite of Argentina. The objectives are monitor both fixed and mobile goods for the transportation industry.
• Rubin-2 (10 kg) a communications satellite of OHB-System, Germany. Rubin-2 uses the ORBComm network comprising 30 satellites to ensure interruption-free communications with the Earth.
• SaudiSat-2 (15 kg) scientific satellite for the Riyadh Space Research Institute, Saudi Arabia
• Trailblazer (110 kg) a technology minisatellite and a mockup of a future commercial lunar orbiter of TransOrbital, La Jolla, CA, USA.
Orbit: Circular orbit, altitude = 650 km, inclination = 65º, period = 97.8 minutes.
RF communications are in the amateur (AMSAT) UHF/VHF bands at 9.6 kbit/s using the AX.25 packet protocol. The UHF transmitting antenna consists of four dipoles providing circular polarization. The VHF receiving antenna is a dipole. The satellite is being operated by the University of Rome.
Mission status: UniSat-2 is no more operational.
AMI (Aerosol Monitoring Instrument)
AMI consists of two wide-angle CCD cameras. One of the cameras is equipped with an INSpector for digital spectroscopy in the spectral range of 430-900 nm. This information is used for precise mapping of the aerosols on Earth surface.
ISIS (In Situ Impact Sensor)
The debris impact sensors are made of piezoelectric devices (analog and digital boards hosting a dedicated microcontroller) used to register the impacts of sub-millimeter particles to obtain statistical information about the LEO environment. 6)
The piezoelectric elements can measure acoustic waves propagating through the satellite structure after an impact of sufficient momentum. Moreover, other effects might be measured in case of direct impacts on the sensors surface (for instance, some piezoelectric materials also show a pyroelectric behavior).
CGM (Cold-Gas Microthruster)
The experimental CGM was jointly developed by Mechatronic System Technik GmbH, Villach, Austria, and by INFM-TASC at Elettra Synchrotron Light Source, Trieste, Italy. The main CGM elements are: control valves and nozzles, pressurized N2 tank, control board, safety valve group, connections, and aluminum plates. The microthruster employs a chamber in which the pressurized nitrogen gas is introduced terminating in a supersonic nozzle in which the gas is accelerated. The throat dimension (0.02 mm diameter) defines the thrust range and is controlled by means of a proportional valve and a pressure sensor. A variable thrust of up to 560 µN is generated. CGM has a total mass of 0.835 kg.
The in-orbit experiment onboard UniSat-2 consists of successive spin-up/spin-down maneuvers using the microthrusters. The satellite nominal spin rate is 3 rpm, with the spin axis aligned to the orbit normal. The gas stored in the tank is enough to perform a sequence of six spin-up/ spin-down maneuvers. One spin-up and one spin-down maneuver per orbit are performed. The time duration of the whole experiment is limited by the microvalve leakage, imposing the experiment to be completed within one day from the safe valve actuation.
The final goal of the CGM project is further miniaturization of the system and the enhancement of its performance for eventual use in a nanosatellite ACS (Attitude Control Subsystem). Such an ACS requires a maximum thrust of about 100 µN that may be achieved using a Hydrazine propellant with an estimated mass flow rate of 0,1 mg/s and a reaction chamber pressure of about 0.3 MPa. A continuous modulation of the thrust is also required; this may be achieved in a control scheme where the pressure in the reaction chamber is used as a feedback signal for controlling the valve opening and thus the mass flow rate. 7) 8)
Legend to Figure 7: 1) microthruster; 2) thrusters group; 3) fluidics connections; 4) safety valve; 5) gas tank; 6) mechanical interface; 7) control unit; 8) power; 9) UniSat-2; and 10) Control Software.
UniSat-3 is the third microsatellite of the University of Rome “La Sapienza”, built by a team of students, researchers and professors at the GAUSS Laboratory.
UniSat-3 is based on the design of the previous two satellites. The spacecraft body is an octagonal prism in shape, 40 cm in diameter and a height of 25 cm. Four aluminium sandwich trays are assembled and kept together by four aluminium columns. Each tray contains a subsystem, allowing for separate manufacturing and testing activities. The trays are octagonal plates of 150 mm side with an overall height of 250 mm for the satellite. The columns are aluminium threaded rods, 6 mm in diameter, encompassing the structure. Eight aluminum panels provide the outer enclosure of the satellite. The solar arrays are surface-mounted on the outside of these panels. 9) 10) 11)
The spacecraft is stabilized using passive magnetic attitude control (no orbit control capability). The system employs a permanent magnet and an energy dissipation system, consists of three magnetic hysteresis rods. The satellite is not equipped with dedicated attitude sensors, except for a COTS three-axes magnetoresistive magnetometer (TAM). Attitude determination (in the order of about 10º) is performed using magnetometer readings and the telemetry data of three solar panels.
An average power of 12.3 W is provided by silicon solar arrays (16% efficiency), an NiCd battery (4 Ah capacity) is used for eclipse operations. The electronic components consist of several boards. The C&DH (Command and Data Handling) subsystem is a compact module, RCM3400, that incorporates a Rabbit 3000 8-bit microprocessor, operating at 29.4 MHz, an 11 bit 8 channel (or 12 bit 4 channel if used in differential mode) A/D converter, 512 kbyte of flash memory, 512 kbyte of static RAM and two clocks (main oscillator and time-keeping). The spacecraft mass is 12 kg.
RF communications: The S/C communications conform to amateur radio standards permitting a data rate of 9.6 kbit/s full duplex communications. The uplink is in VHF (145 MHz), the downlink in UHF (436 MHz). The communications protocol used is the HDLC (High-level Data Link Control). Each piece of raw data is encapsulated by the microprocessor in an HDLC frame by adding a trailer and a header.
A launch of UniSat-3 took place on June 29, 2004, on a Dnepr launch vehicle from the Baikonur Cosmodrome, Kazakhstan.
UniSat-3 was part of a multiple payload launch consisting of the following additional satellites: SaudiComsat-1, SaudiComsat-2 and SaudiSat-2 from Saudi Arabia (KACST); AprizeSat-1 (LatinSat-C), AprizeSat-2 (LatinSat-D), and AMSAT OE (Echo) from USA (SpaceQuest); DEMETER from France (CNES), AKS-1 from Russia (Aerospace Systems), and Celestis-04, a space burial from USA (CPAC).
Orbit: Sun-synchronous orbit, mean altitude = 750 km, inclination = 98º. Local equator crossing time on ascending node at 22:30 hours.
UniSat-3 is operational in 2010 (more than 5 1/2 years after launch). The data of the spacecraft are being received at the University of Rome ground station on a regularly basis.
• The magnetometer experienced a z-axis failure during the first year in orbit. 12)
All the electronic boards and photovoltaic system can be considered as payloads, as they are realized with commercial electronic components, not qualified for space environment. 13) 14)
Solar Cell Experiment (Power Subsystem)
Four experimental solar panels are installed on the lateral surfaces in order to test different solutions. Several bonding techniques were employed.
• Terrestrial grade monocrystalline silicon solar cells (BP-Saturn7), encapsulated using an improved terrestrial technology procedure, are the basis of the photovoltaic system.
• One panel was made by Kiev Polytechnic Institute of the National Technical University of Ukraine. This panel was manufactured with monocrystalline silicon solar cells, covered with glass. The nominal efficiency is about 15%. The cells are bonded to an aluminum honeycomb panel using polyimide film as dielectric insulation.
• Three panels were manufactured using triple junction GaAs cells with different cover materials to test in-orbit degradation. The nominal efficiency of the triple junction solar cells is 21%. Each array is realized by two parallel arrays. Each string has a series of four solar cells (Figure 15).
The power subsystem features a newly developed MPPT (Maximum Peak Power Tracking) device; the objective is to track the solar panels in such a fashion to obtain the maximum power operative condition, adapting to changing illumination and S/C temperature. The MPPT is based on a step-up DC/DC converter, in which the battery voltage is higher than the solar array voltage. The solar array connected to the MPPT is made of 14 silicon solar cells, with 7.6 V nominal open circuit voltage.
Magnetometer (Honeywell HM2003) for the measurement of the Earth's magnetic field (strength and direction). The 3-axis device is mounted on a PCB (Printed Circuit Board) containing the management circuitry.
MPPT (Maximum Peak Power Tracking)
The prototype MPPT system has been designed, manufactured and tested to be demonstrated on the satellite. The objective of MPPT was to improve the performance of the onboard power generation system. The MPPT includes an adjustable DC/DC converter connecting the solar panel to the load. In this way the system is capable to track autonomously the solar array maximum power operative condition, adapting to changing illumination and temperature events during the orbital phases. The MPPT configuration implemented is considered an upgrade configuration to conventional systems, in which the voltage of the battery is higher than the voltage of the solar arrays.
The MPPT has been manufactured on a PCB (Printed Circuit Board) consisting of six layers. SMD (Surface Mount Device) components have been exploited to reduce the volume (50 mm x 82,4 mm) and the mass of the circuitry to ~ 50 g. All other electronics of the system are COTS (Commercial-of-the-Shelf) components. 15)