Concept of Operations ConOps for LEISA CubeSat Nicholas Clegg, Scott Rossol 7/28/11 This document describes the characteristics of the program known as LEISA (Low Earth Ionospheric Spectrum Analyzer). This program is a project of COSMIAC (Configurable Space Microsystems Innovations & Applications Center) a division of UNM (University of New Mexico) who is funded by NSF (National Science Foundation), AFRL (Air Force Research Laboratories), PTI (Philips Technology Institute), SNL (Sandia National Labs), and LANL (Los Alamos National Labs). 2 Concept of Operations COSMIAC Contents 1. Overview 1.1 Mission Identification 1.2 Document overview 1.3 System overview 2. Current System Situation 2.1 Mission Environment 2.1.1 The Thermosphere 2.1.2 Ionosphere 2.1.3 Plasma 2.1.4 Effects of Plasma 2.1.5 Thermal Environment 2.1.6 Radiation Environment 2.1.7 Orbital Analysis 2.1.8 Access Time 2.1.9 Global Education Network for Satellite Operations (GENSO) 3. Launch Opportunities 4. System Budget 4.1 Projected Budget 5. System Components 5.1 1U CubeSat Structure 4.1.1 COSMIAC Wiki Page 5.2 Deployable Antenna System 5.2.1 Antenna 5.2.2 Specifications 5.2.3 Antenna Configurations 5.2.4 Features 5.3 Main Antenna System 5.3.1 Main Antenna 5.4 NanoPower Solar 100U 5.4.1 Overview 5.4.2 Features 5.4.3 Solar String 5.4.4 Sun Sensor 5.4.5 Temperature Sensor 5.4.6 Gyroscope 5.4.7 Magnetorquer 3 5.4.8 Product Properties 5.5 Electrical Power System 5.5.1 1U CubeSat EPS 5.5.2 Features 5.5.3 Dimensions 5.5.4 Electrical Characteristics 5.5.5 Battery Charge Conditions 5.5.6 Battery Discharge Conditions 5.5.7 Unregulated Battery Bus 5.5.8 Communications 5.6 Texas Instruments ADC 5.6.1 Basic Functions 5.6.2 Input 5.6.3 Output 5.6.4 Power 5.6.5 Frequencies 5.7 Actel FPGA Board 5.7.1 Overview 5.7.2 Connectors 5.7.3 Output 5.7.4 Power 5.8 Communication System 5.9 AstroDev He – 100 5.9.1 Over View 5.9.2 Dimensions 5.9.3 Frequencies 5.9.4 Max Data Rate 5.9.5 Receiver Sensitivity 5.9.6 Power 5.9.7 Operating Temperature 5.10 VHF uplink Receiver 5.10.1 Dimensions 5.10.2 Specifications 5.10.3 Power 5.11 NanoCom Half-Duplex Transceiver 5.11.1 Dimensions 5.11.2 Frequencies 5.11.3 Specification 5.11.4 Power 5.12 Oem Industrial Wireless Modem 4 5.12.1 Dimensions 5.12.2 Specifications 5.12.3 Sensitivity 5.12.4 Power 5.13 Lightning Detection Antenna 5.13.1 Overview 5.13.2 Antenna Types 5.14 GPS 5.15 San Jose Technology, Inc 5.15.1 Dimensions 5.15.2 Connector 5.15.3 Output 5.15.4 Start times 5.15.5 Specifications 5.15.6 Power 5.16 GlobalSat 5.16.1 Dimensions 5.16.2 Connectors 5.16.3 Start Times 5.16.4 Power 5.16.5 Specifications 5.17 Locosys 5.17.1 Specifications 5.17.2 Power 5.18 Passive Magnetic Attitude Stabilization System 5.18.1 Stabilization System 6. Ground System 6.1 Introduction 6.2 Transceiver 6.2.1 Icom IC-910H 6.2.2 Dimensions 6.2.3 Power Supply 6.2.4 Specifications 6.2.5 Frequency Coverage 6.2.6 Receiver 6.2.7 Spurious Rejection Ratio 6.2.8 Transmitter 6.3 CubeSat Launch History 7. Operational constraints 7.1 Budgetary Concerns 5 7.2 Research Time Constraints 7.3 Launch Availability Constraints 8. Support environment 9. Operational Scenarios 9.1 Possible Scenarios 10. Notes 11. Appendices 12. Acronym Table 6 1. Overview 1.1 Mission Identification The mission of the project is to help understand the effects of the ionosphere on the propagation of electromagnetic (EM) waves. Fluctuations in the electron content of the ionospheric layers are the principal phenomena that affect EM wave propagation. By studying the propagation of waves through the ionosphere, we hope to understand these phenomena. Phase 1: To produce a ground-based system to detect radio frequency (RF) emissions from lightning within the frequency band of 30-300MHz. Phase 2: To create three second-revision ground-based systems that will be installed around the state and will enable the triangulation of lightning. Phase 3: To integrate the ground detection system into a compact low-power CubeSat form factor. 1.2 Document overview To establish a general document used to describe the characteristics of this operation, its constraints and its goals. This document will include statements of goals that are set for the operation, strategies, tactics, policies, and constraints affecting the system. 1.3 System overview 1.3.1 1.3.2 1.3.3 1.3.4 1.3.5 1.3.6 1.3.7 1.3.8 1.3.9 Downlink? (trailblazer – 0.2-3w UHF @ 9600bps) Uplink? (trailblazer VHF) Orbit – 320km at 60⁰ Inclination Battery system – Clyde-Space 20Whr Electrical Power System (EPS) – Clyde-Space C&DH module – Texas Instruments ADC Actel – FPGA board for storage GPS Unit Mission to generate spectrograms of lightning chirps 7 2. Current system situation 2.1 Mission environment 2.1.1 The thermosphere is the largest section of Earth’s atmosphere. It spans from 80 km to 500-1,000 km above the surface of the earth. Due to this high levels of solar radiation and the large spacing of molecules in the thermosphere temperatures of the molecules can reach around 2000°C and can fluctuate by hundreds of degrees depending on the quantity of solar radiation. These same properties also make the temperature in the thermosphere increases with altitude. These great ranges in temperature that are driven by the sun cause periodic oscillations in the thermosphere accounting for the large range in the thermosphere’s upper bound. Theses oscillations are compared to tides on Earth just affected by the solar day and not the position of the moon. Also, as the few molecules in the thermosphere are bombarded with relative high doses of radiation they ionize. This is why the Thermosphere is part of the ionosphere. Even though the thermosphere is a part of earth’s atmosphere its density is so low that a majority of it is what we typically think of as space. 2.1.2 The Ionosphere The ionosphere spans an altitude from 50 km to 1000 km in the Earth’s atmosphere. It consists of parts of the exosphere, thermosphere, and the mesosphere. The ionosphere is defined by the fact that in this region is ionized by solar radiation, primarily UV, X-Ray, and other shorter wavelength Electromagnetic Radiation. The solar radiation ionizes the atmosphere with a power density of 1370 W/m^2. These ionized particles create plasma in this region that can have great effects on radio signals. The ionosphere is made up of multiple layers. These layers were first purposed by Oliver Heaviside in the early 1900’s. The first layer is the D layer. It is the lowest layer of the ionosphere spanning from 60 -90 km being the lowest in the atmosphere the D layer is the least ionized and has very little effect on radio signals, only to reduce the strength of high frequency radio signals. The next layer is the E layer which excises from 90 to 120 km. The E layer is a little bit more ionized then the D layer and therefore can reflect radio waves but normally only those below 10 Hz. However, sporadic E-layers can form in the E-layer and have the ability to reflect radio waves to 50 Hz and higher, sometimes close to 225 Mhz. Sporadic E-layers are defined as small but intense clouds of ionized particles that appear in the E-layer. The final layer is the F layer. That spans from 200 km to 500 km. If radio waves penetrate the Flayer from earth they will most likely reach open space, however the F-layer is reflects the most radio waves and is the only layer with significant ionization at night, and in the day increases in ionization it splits into two layers the F1 and F2 regions. F1 only exists in day time where F2 is sustained through night time. 2.1.3 The Thermosphere Plasma Plasma is known as the fourth state of matter. It is what makes up all stars in the known universe and much of the space surrounding them. Plasma has no shape definite shape or volume, like gas, but plasma is unique in the fact it is affected by electromagnetic fields. This is because plasma is formed when elements, normally in gaseous phase, are ionized through the introduction of heat or 8 another energy source such as radiation. Ionization is when an element gains or losses an electron. In plasma a group of atoms are gaining or losing electrons, which in other terms means that electrons become free to move between the atoms in the plasma. This movement of electron is what makes plasma electrically conductive as well as susceptible to electromagnetic fields. A distinguishing factor of plasma is that the electrons of the molecules in plasma interact not only with the closes molecule but multiple molecules in a given area around it. 2.1.4 Effects of Plasma In polar orbits or at high altitudes plasma environments can electrostatic charge the space craft surfaces. At low altitudes plasma can cause power leakage from exposed solar arrays and electromagnetic perturbations. At low altitudes surface charging is not a problem but because the satellite travels at a mesosonic speed, faster than the ion thermal speed and slower than the electron thermal speed a wake is formed behind the spacecraft as the ions do not have the time to repopulate the area and surfaces and bodies in this area can become charged if exposed to energetic electron fluxes. Low altitudes also allow systems to drive current through local plasma if operating at elevated voltages. The event in which currents become larger than expected because of plasma sheath geometry is called snap over. Ionospheric plasma, which exists in low orbits, is in nature non-homogenous and dispersive. This causes interactions with electromagnetic equipment like telecommunication devices, global positioning systems and RADAR. 2.1.5 2.1.6 2.1.7 Thermal Environment Radiation Environment Orbital analysis Altitude 300 – 400 km Inclination 60 degrees Expected orbital life time is ~50 days Orbit data Orbital Data Based on orbit altitude of 320 km or 198.8 miles Data Value Units Maximum Angular Altitude 1,271.325 Miles Maximum Angular Altitude 2,046.000 km Maximum Angular Range 2,542.651 Miles Maximum Angular Range 2,046.000 km Maximum Surface Range 2,433.290 Miles Maximum Surface Range 3,916.000 km Satellite Orbit Velocity 4.770 Miles/Second Satellite Orbit Velocity 17,269.148 Miles/Hour Satellite Orbit Velocity 7.720 km/Second Satellite Orbit Velocity 27,792.000 km/Hour Orbit Revolution Time 90.000 Minutes Visibilty Half-Angle 17.800 Degrees Maximum Visibility Tme 8.900 Minutes 9 2.1.8 Access Time On average 12 incidents of access every 48 hours Access 1 2 3 4 5 6 7 8 9 10 2.1.9 Start Time Stop Time Duration 38:06.2 47:18.5 552.237 13:05.4 20:31.4 445.983 36:13.3 44:13.4 480.099 09:42.6 18:40.0 537.365 22:02.6 28:17.8 375.214 53:59.3 03:16.4 557.161 30:55.2 35:28.5 273.329 20:02.7 23:17.0 194.319 51:47.4 00:56.0 548.614 26:21.1 33:29.6 428.452 Average--> 439.2773 Global Education Network for Satellite Operations (GENSO) 3. System Budget 3.1 Projected budget 10 4. Launch Opportunities 4.1 ISIS Launches Launch Opportunities Orbit Inclination 190 x 380 km 52° 450 - 500 km Sun Sync 2012 450 - 550 km 79° 600 - 700 km Sun Sync 600 -700 km Sun Sync 540 / 820 km Sun Sync 325 x 1500 km 80° 450 - 500 km 65° 2013 450 - 550 km 79° 450 - 500 km 98° 600 - 800 km Sun Sync 620 km Sun Sync 620 km Sun Sync 600-650 km Sun Sync 2014 450 - 550 km 79° Geo Sync Geo Sync Low Lunar Low Lunar 2011 5. System Components 5.1 1U CubeSat Structure 4.1.1 Link to Wiki Structure 5.2 Deployable Antenna System 5.2.1 5.2.2 Antenna Contains up to four tape spring antennas of up to 55 cm length Accommodate up to four monopole antennas, which deploy from the system after orbit insertion Able to configure antenna system to be compatible with all UHF and VHF radios typical of CubeSat systems Specifications RF impedance: 50 Ohm Max RF power: 2 W Insertion loss: 1.5 dB Frequency range: 130 – 165 MHz (VHF), 390 – 450 MHz (UHF) Electrical power: nominal/standby < 20 mW 11 5.2.3 5.2.4 Electrical power: during deployment 2 W Total weight less than 100g depending on antenna configuration Envelope stowed: L x W x H = 98mm x 98mm x 7mm Operational temperature range: -30 to 70 C The following antenna configurations are possible in UHF and/or VHF 1 Turnstile 2 Dipoles 1 Dipole and 1 Monopole 1 Dipole and 2 Monopoles 1/2/3/4 Monopole Features Antenna polarization circular or linear Deployment confirmation switch per antenna Software safe/arm implementation Dual redundant deployment system Compatible with ISIS products and recent Pumpkin, Clyde Space and Gom Space products 5.3 Main Antenna System 5.3.1 5.3.2 Main Antenna 5.4 NanoPower Solar 100U 5.4.1 Overview The NanoPower Solar 100U is an integrated side panel solution for CubeSat satellites which provides solar cells, Magnetorquer, sun sensor, gyroscope and temperature sensor all on a single PCB only 2mm thick. The solar cells are space qualified triple junction cells from AzurSpace giving an efficiency of 28% or more. Two cells are reflow soldered to the panel and connected in series to achieve an output voltage of approximately 4.6V suitable for the NanoPower P-series power supply systems from GomSpace. A Magnetorquer is integrated into the PCB in ten internal layers thus taking up no extra space inside the spacecraft, and with an effective area of 1.6 m2 it is by far enough to stabilize and control the attitude of a CubeSat. To assist in attitude determination, the panel also features a small photo diode intended to be used as a sun directional sensor. A temperature sensor is mounted close to the diode to enable compensation for its temperature drift. In addition, a MEMS gyroscope is mounted on the inward-facing side of the panel giving a rotational resolution of 0.004 o/s. The attitude determination electronics on the panel interfaces seamlessly to the NanoMind A712 on-board computer via a single connector including power supply and SPI bus for temperature sensor and gyro, output from sun sensor. The Magnetorquer uses its own connector allowing high switching currents, and the solar cells also use a separate connector to connect to the power supply system. 12 5.4.2 5.4.3 5.4.4 5.4.5 5.4.6 5.4.7 5.4.8 Features Two series-connected AzurSpace 3G-28 space qualified triple junction solar cells and protection diodes Integrated Magnetorquer of 1.6 𝑚2 Coarse sun sensor Temperature sensor Gyro-scope model ADIS16251 (0.004 o/s) Solar Cell String (Condition: Full sunlight in LEO) Voltage: 4.64 to 4.84V Current: 490 to 508 mA Power: 2270 to 2400 mW Efficiency: 27.7 to 29.1% Sun Sensor Current : 170 uA Temperature Sensor Range: -55 to 150°C Resolution: 1.5 to 3.5°C Vcc: 3.3 V Current: 260 uA typ. Temperature coefficient:0.233 %⁄℃ Cosine error: 1.85° typ. Gyroscope Range: 80 °⁄s Sensitivity: 0.00458 °⁄s Bias Stability: 0.016 °⁄s Vcc: 5 V Current 44 mA Magnetorquer Area: 1.6 𝑚2 Resistance: 106 to 130 Ohm Current(Absolute Maximum Rating): 0.7A Dipole Momentum at 3.3V: 0.04 to 0.05 𝐴𝑚2 Product Properties Material: Fr4 Tg 180 Board Operational temperature: -40 C to +85 C Dimensions: 82.5mm x 98.0mm x 2.1mm Height of connectors: 4.8 mm Height of gyro: 5.5 mm Panel Thickness: 2.15 mm Mass: 59g 13 5.5 Electrical Power System 5.5.1 5.5.2 5.5.3 5.5.4 5.5.5 5.5.6 5.5.7 5.5.8 1U CubeSat EPS with 20 watt hour integrated battery The batteries utilize Lithium Ion Polymer technology to offer world leading power to mass ratios in a form factor ideally suited to the volume constraints of CubeSats. In addition to this, testing has been carried out by both ESA and NASA, and the batteries have been cleared for launch on NASA manned flights. Features Integrated active solar array maximum power point tracking Battery charge management for lithium ion and lithium polymer batteries Battery under-voltage protection Raw batter buses with over-current protection Telemetry and telecomm and via I2C True dead launch – ideal diode and separation switch USB battery charger for ground testing Dimensions (l x w x h) With 20 Whr Batteries : 95mm x 90mm x22mm With 20 Whr Batteries: 229 g Electrical Characteristics Input voltage: 3.5 – 8 V Output voltage: 6.2 – 8.26 V Output current: 0 – 0.5 A Operating frequency: 160 – 180 KHz Efficiency at 6V Input, Full Load: 77% - 80% Batter Charge Conditions EoC Voltage: 8.22 – 8.30 V Charge Current: 0.625 A Battery Discharge Conditions Full Discharge Voltage: 6.16 – 6.24 V Discharge Current: 1.25 A Recommended Depth of Discharge: 20% Unregulated Battery Bus Output Voltage: 4.95 – 5.05 V Output Current: 4 A Efficiency at 8.62 V Input, Full Load 98.5% - 99.5% Communications Protocol: 𝐼 2 𝐶 Transmission Speed: 100 – 400 KBps Bus Voltage 3.26 – 3.33 V Address Scheme: 7 bit Node Operating Frequency: 8MHz 14 5.6 Texas Instruments ADC 5.6.1 5.6.2 5.6.3 5.6.4 Basic Functions Two analog inputs to the ADC are provided via external SMA connectors. One input path uses a pair of THS9001 amplifiers, while the other input is ac-coupled. In both cases, the user supplies a single-ended input, which is converted into a differential signal. The EVM provides an external SMA connector for input of the ADC clock. The single-ended input is converted into a differential signal at the input of the device. Digital output from the EVM is via a high-speed, high-density Samtec output header. A breakout board is provided for high-speed logic analyzer touchless probing solutions from both Agilent and Tektronix. Both the Agilent E5405A and Tektronix P6980 probes are supported. Power connections to the EVM are via banana jack sockets. Separate sockets are provided for the ADC analog and digital supplies and for the differential amplifier supply Input 2 Analog Inputs from external SMA connectors 1 external SMA connector from ADC clock Output Samtec output header Supported by Agilent E5405A and Tektronix P6980 Connected to Breakout Board Power Separate hookups for ADC analog and digital supplies and for both amplifiers Voltages must be met or damage or false processing may occur Three types of power supplies: 5.6.5 3.0 – 3.6V ADC analog driver supply 3.0 – 3.6V ADC digital driver supply 4.75 – 5.25V ADC analog supply Frequencies Clock frequency should not exceed 500 MHz Default board forms a bandwidth filter with a pass bandwidth of 4 MHz and 1 GHz 15 5.7 Actel FPGA Board 5.7.1 5.7.2 5.7.3 5.7.4 Overview The Cortex-M1 Enabled IGLOO TM Development Kit is an advanced microprocessor based FPGA development and evaluation kit. The purpose of the kit is to help the user become familiar with the IGLOO FPGA features by providing a useful Sample Design, with a “How To” tutorial for implementing the FPGA hardware design using Libero Project Manager and CoreConsole. Connectors USB 20 – pin Cortex – M1 JTAG connector LVDS Connectors GPIO Connectors Expansion Output Desired output 200 MHz Clock Frequency is 48 MHz Power Core runs from 1.2 – 1.5 V Input from 5V 2.1 mm positive center power supply Draws 100mA to 500mA of 5V current to consume 500mW to 2500mW 5.8 Communications System 5.9 AstroDev He – 100 5.9.1 5.9.2 5.9.3 5.9.4 5.9.5 5.9.6 Over View Provides a CubeSat kit compatible communication system for extreme environment applications. Dimensions 90.17mm x 96mm x 1.63mm 78g Frequencies Tx: 120 – 150 or 400 – 450 MHz Rx: 400 – 450 or 120 – 150 MHz Max Data Rate 38.4 kbps Higher speeds under tests Receiver Sensitivity -104.7 dBm at BER 10−3 Power Input Voltage: Logic – 3.3 V Input Voltage: Transmitter 5 – 16 V Output Transmitter Power: 100 mw – 3 W 16 5.9.7 Power Usage: Receiver - <200 mw Power Usage: Transmitter - <6 W Operating Temperature -30 C to +70 C 5.10 VHF Uplink Receiver 5.10.1 Dimensions 90 mm x 96 mm x 5 mm 60 g 5.10.2 Specifications Frequencies: 130 – 160 MHz Max Data Rate: 1.2 kbps Receiver Sensitivity: -100 dBm at BER 10−5 Operating temp: -20 C to +60 C 5.10.3 Power Input Voltage: 3.3 V Power Usage: < 0.2 W 5.11 NanoCom Half-Duplex Transceiver 5.11.1 Dimensions 90 mm x 96 mm x 11 mm 75g 5.11.2 Frequencies Rx: 432 – 438 MHz Tx: 432 – 438 MHz 5.11.3 Specifications Receiver Sensitivity: -115 dBm for 1200 bans detection threshold Max Data Rate: 400 kbps Operating Temperature: -30 C to +60 C 5.11.4 Power Input Voltage: 3.3 V Output power: 200 – 600 mW Spurious Transmission: 0.025uW 5.12 Oem Industrial Wireless Modem 5.12.1 Dimensions 89 mm x 53.4 mm x 17.8 mm 55 g 5.12.2 Specifications Frequency: 402 – 428 MHz Max Data Rate: 230 kbps Operating Temperature: -40 C to +85 C 5.12.3 Sensitivity -116 dBm at 9.2 kbps -108 dBm at 172 kbps 5.12.4 Power 17 Input Voltage: 3.3 to 5.5 V Output Power: 100 mW to 1 W 5.13 Lightning Detection Antenna 5.13.1 Overview The lightning detection antenna on LEISA has three primary design requirements: it must be deployable from a 1 U cubesat, it must be circularly polarized, and it must be able to detect RF signals between 20 to 300 MHZ (has a bandwidth of over 10:1). There are relatively few antennas that fit any of these parameters, so that finding an antenna that fits all three requirements is a major challenge. This problem is made more difficult by the fact that the smaller the operating frequency than the larger the antenna, making a 20 MHZ antenna inherently difficult to deploy from a 10 cm cube. The process for designing the antenna consists of first considering known antenna types and selecting one that best fits the requirements, then using HFSS to design the dimensions and simulate the antenna. 5.13.2 Antenna Types Antenna types considered, and reasons rejected: Dipole antenna: Narrow band Linear polarization Helix antenna: Narrow band Horn antenna: Difficult to deploy Yagi: Narrow band Linear polarization Difficult to deploy Log Periodic: Linear polarization Difficult to deploy Fractal antenna: Not enough information The two current antennas that are being considered are biconical antennas and spiral antennas. There are biconical antennas that have already been designed for the necessary bandwidth, and we have a working idea on how to deploy one from a CubeSat satellite. However, the biconical antenna is linearly polarized, it could be made to be circularly polarized by deploying two antennas perpendicular to each other; but this makes them significantly harder to deploy. Spiral antennas are broadband and circularly polarized; however there is little information on how to design 18 one that will operate within the required frequency range and a spiral antenna would be difficult to deploy from a CubeSat. 5.14 GPS 5.15 San Jose Technology, Inc 5.15.1 5.15.2 5.15.3 5.15.4 5.15.5 5.15.6 FV- M8 ($99.95) Full EMI Shielding Dimensions 30mm x 30mm x 8.6mm 15g Connector 8 pin with 1.0 mm pitch Output RMC GGA GSV VTG GSA GLL GDA Start times 1-41 sec Specifications Sample rate: 1-5 Hz Accuracy: ~3m Sensitivity: -158 db Power 3.3 – 5 V at 33- 63 mA 315-165 mW 5.16 GlobalSat EM-408 ($65) Extra RF shielding Extra Antenna 5.16.1 Dimensions 36.4mm x 35.4mm x 8.3mm 20 g 5.16.2 Connectors 5-pin interface cable NMEA 0183 GGA GSA 19 GSV RMC VTG GLL 5.16.3 Start times 8-42 sec 5.16.4 Power 25-75mA at 3.3 V 82.5 – 247.5 mW 5.16.5 Specifications Sample rate? Probably ~1Hz Accuracy: ~10m Sensitivity: 159db Max Altitude: 18km Max Velocity: 515m/s 5.17 Locosys LS20030~3 ($59.99) Built in battery 5.17.1 Specifications 16mm x 35mm x 6.8mm Accuracy: ~3m Sensitivity:? 5.17.2 Power 29 – 32 mA at 3.6 V 104.4 – 115.2 mW 5.18 Passive Magnetic Attitude Stabilization System 5.18.1 Passive Magnetic Attitude Stabilization System Trailblazer is implementing one Simple and reliable Cheapest solution for attitude control Stacking height is less then 15mm Thermal range -50 C to +100 C Mass 60 to 90g 6. Ground System 6.1 Introduction 6.1.1 Consists of two circularly polar radar towers one tuned to VHF (144.0-148.0 MHz) and the other to UHF (440.0-450.0 MHz). Towers transfer data through two Terminal Node Controllers (TNCs) one able to transfer 1200 bps the other up to 9600 bps to computer banks. The system runs of the ICOM 910 Transceiver. 6.2 Transceiver 6.2.1 Icom IC-910H 20 http://www.icomamerica.com/en/products/910h/specifications.aspx http://www.icomamerica.com/en/products/910h/ 6.2.2 6.2.3 6.2.4 6.2.5 6.2.6 Physical 241 mm x 84 mm x 239 mm (WxHxD) 4.5kf: 10lb Power Supply 13.8 V DC + 15% (negative ground) Specifications Data: 9600 bps Packet Operation Stability: + 3 ppm (-10 to 60 C) Resolution: 1 Hz Frequency Coverage Tx: VHF: 144.0-148.0 MHz UHF: 440.0-450.0 MHz Rx: VHF: 132.0-174.0 MHz UHF: 440.0-480.0 MHz Receiver VHF SSB, CW FM UHF SSB, CW FM Single Conversion Super Heterodyne Double Super Heterodyne Double Super Heterodyne Triple Super Heterodyne MAIN band SUB band Mode 144 MHz band 430 MHz band 1200 MHz band*1 1st 2st 3st 1st 2st 3st SSB 10.8500 - - 10.9500 - - CW 10.8491 - - 10.9491 - - FM 10.8500 0.4550 - 10.9500 0.4550 - SSB 71.2500 10.8500 - 71.3500 10.9500 - CW 71.2491 10.8491 - 71.3491 10.9491 - FM 71.2500 10.8500 0.4550 71.3500 10.9500 0.4550 SSB 243.8500 10.8500 - 243.9500 10.9500 - CW 243.8491 10.8491 - 243.9491 10.9491 - FM 243.8500 10.8500 0.4550 243.9500 10.9500 0.4550 21 Receiver System Sensitivity (µV) Squelch sensitivity (µV) Selectivity SSB, CW 0.11 at 10 dB 1.00 at threshold 2.8 (KHz/-6dB)<x<4.2 (KHz/-60dB) FM 0.18 at dB 0.18 at threshold 15(KHz/-6dB)<x<30(KHz/-60dB) FM-N NA NA 6(KHz/-6dB)<x<18(KHz/-60dB) 6.2.7 Spurious Rejection Ratio <60 dB (Except at 1200M Hz Bad) RIT Variable Range SSB, SW 1.0 kHz (2.0 kHz at 1200 MHz band) FM 5.0 kHz (10.0 kHz at 1200 MHz band) 6.2.8 Transmitter Frequency 144 MHz 440 MHz 1200 MHz 6.3 Output power 5-100 W 5-75 W 1-10 W Spurious emission >-60 dB >-60 dB >-50 dB CubeSat Launch History CubeSats were developed in the turn of the century by California Polytechnic State University. The first major launch of Cube satellites was in 2003 where 6 CubeSats were launch in Eurockot LV from Plesetsk, Russia. CubeSat Launches continued to launch from rockets out of Russia, Kazakhstan, India, and a long with larger shuttles such as the Endeavour in 2009. From 2003 to 2009 a total of 33 1U, 3 2U, and 6 3U CubeSats were launched. Ford details on the launches and the specific CubeSats launched see Michael’s List of CubeSat Satellite Missions http://mtech.dk/thomsen/space/cubesat.php 7. Operational Constraints 7.1 Budgetary Concerns 7.2 Research Time Constraints 7.3 Launch Availability Constraints 8. Support Environment 22 9. Operational Scenarios 9.1 Possible Scenarios 10. Notes 11. Appendices 23 12. Acronym Table Acronym Definition ADC BER C&DH CW EM EMI EPS ESA EVM FM FPGA GDA GENSO GGA GLL GPIO GPS Analog-to-Digital Converter Bit Error Ratio Command & Data Handling Continuous Wave Electromagnetic Electromagnetic Interference Electrical Power System European Space Agency Evaluation Module Frequency Modulation Field-Programmable Gate Array Generic Data Access Global Educational Network for Satellite Operations Generalized Gradient Approximation Geographic Latitude and Longitude General Purpose Input/Output Global Positioning System GPS receiver operating mode, satellites used in the position solution, and DOP value The number of GPS satellites in view, satellite ID numbers, elevation, and azimuth Low Earth Ionospheric Spectrum Analyzer Low Earth Orbit Low Voltage Differential Signal Micro Electro-Mechanical Systems National Aeronautics and Space Administration Nano Satellite Sensor Package Radio Detection And Ranging Radio Frequency Receive Incremental Tuning Time, date, position, course and speed data Receiver Subminiature version A Coaxial RF Single Side Band Terminal Node Controllers Transmitter Ultra High Frequency Universal Serial Bus Ultra Violet Very High Frequency GSA GSV LEISA LEO LVDS MEMS NASA NaSSP RADAR RF RIT RMC RX SMA SSB TNCs TX UHF USB UV VHF 24 VTG Whr Course and Speed Information Relative to Ground Watt Hour