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 2
 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 2
 Resistance: 106 to 130 Ohm
 Current(Absolute Maximum Rating): 0.7A
 Dipole Momentum at 3.3V: 0.04 to 0.05 2
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:  2 
 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
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