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This is the slide show from our preliminary design review for our Junior Engineering Project.
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Preliminary Design ReviewTeam ESAT
Taylor UniversityJunior Engineering ProjectPicoSatellite
Caleb CarrollMarc CattrellElliot ChalfantLuke DornonZach Vander LaanDavid Zilz
Advisors:Dr. Hank VossMr. Jeff Dailey
1
What is a PicoSatellite?◦ Satellite some where between 0.1-1 kg◦ Often flown in groups of 3 or more
How Does is differ from a regular satellite?◦ Lower Orbit◦ Able to reach unexplored areas of Atmosphere◦ Lower Cost
Introduction
2
TubeSat◦ Able to fit into tube for InterOrbital Flight◦ Deployable from tube ready for orbit◦ Withstand launch conditions
CubeSat◦ Constellation of PicoSatellites◦ Fit 4-5 smaller satellites into CubeSat
Project Scope
3
Why are we doing this?◦ Small Size/Weight/Cost◦ Modular
(Compatibility w/ other system)
◦ Stability Control◦ Power Management◦ Solution for Thermal
Problems◦ Scientific Instruments
Project Requirement
4
Work Breakdown StructureESAT 1.0
Taylor UniversityJunior Engineering
Project1.1Manage
mentMarc
1.1.1 Project
Statement
1.1.2 Project
Requirements
1.1.3 Specification Sheet
1.1.4 Gantt Chart
1.2Mechanic
alCaleb
1.2.1Enclosur
e
1.2.2Deploym
ent
1.2.3Thermal
1.2.4Aerodyna
mics
1.3Comm.
Prof. Dailey
1.4 Micro
processor
Prof Dailey
1.5 Power
Management Luke & David1.5.1 Power
Supplies
1.5.2 Power Needs
1.6 Sensors
Marc
1.6.1 Plasma Probe
1.6.2 Magneto
meter
1.6.4 Electrical
Schematics
1.6.5 Data
Collection
1.7Attitude ControlZach & Elliot1.7.1
Preliminary Research
1.7.2 Engineering Requirement
s
1.7.3 Visual
Representations
1.7.4 Magnet Testing
1.8Testing
All
1.8.1Circuitry Testing
1.8.2Stress Testing
1.8.3Comm. Testing
1.8.4Sensor Testing
5
Block Diagram
6
Category Component Qty Mass (g) Voltage (V) Current (mA) Power (W) Duty cycle Total mass (g) Total power (W)
Power supply Solar panels 12 6.5 2max 400 1 78 0.8
Power supply Batteries 4 875 0 0
Power supply Power regulator 1 1.5 1 0 1.5
Control Command interface 1 3.3? 10 1 0 0.033
GPS GPS unit 1 3.5 3.3 rated 3-5 30 1 3.5 0.099
Communication MaxStream 1 18 1 18 0
Communication Satellite transmitter 1 0.12 0 0 0
1 2.5 1 0 2.5
Measurement Magnetometer 1 4 6.5 20 0.013 1 4 0.013
Measurement Plasma probe 2 50 6.5*+/- 2.5 0.0325 1 140 0.065
Total mass 243.5g
Total power supply 0.8W
Total power use 4.21W
System Requirements
7
GOAL: Stabilize satellite and meet attitude control objectives using magnetic stabilization
What are attitude control issues to consider?• Antenna orientation• Sensor orientation (Plasma Probe, Magnetometer)• Power constraints (Solar Panels)• Aerodynamics
1.7 Attitude Control
8
Permanent Magnet 2 Permanent Magnets (perpendicular
orientation) Motor-controlled Magnet Permanent Magnet with Magnetic Torquer Magnetic Torquers for 3 axes Reaction Wheels / Thrusters Gravity Gradient Boom
Attitude Control - Options
9
Solar panels on both sides of satellite◦ From power standpoint do not need to control
satellite’s roll Advice from Taylor Engineering alum
◦ Strongly urged us to scale back scope of project◦ Simple magnetic stabilization would be
sufficiently difficult for semester - long project
Use permanent magnet as method of attitude control
Attitude Control – Narrowing the Scope
10
Attitude Control – How Does a Permanent Magnet Work?
[1]
[2]
Earth modeled as a dipole magnet with roughly 11 degree angle of declination from geographical poles.
Magnetic torque due to interaction between permanent magnet and Earth’s magnetic field
Need magnetic torque to be greater than any other torque on satellite• Satellite will track the magnetic field of the
Earth, rotating twice per orbit.
𝜏= 𝜇Ԧ × 𝐵ሬԦ
11
At orbit of r=310km and T=1000K:◦ For altitudes below 500km, drag force dominates all other forces (such as
radiation) [3]
Permanent magnet controls 2 axes (pitch and yaw) but roll appears to be unconstrained. ◦ While satellite may be able to roll over equator, it will not be able to do so near
the poles Drag force constrains the 3rd axis
Two surfaces of satellite will be in Ram direction during different parts of orbit ◦ These surfaces due to drag force are working against magnetic torque◦ How large are these torques in a worst case scenario?
Attitude Control – The Drag Force Drag Force= .02 𝑑𝑦𝑛𝑒𝑠/𝑐𝑚2 [3]
12
Worst case scenario for torque due to drag force: ◦ 10’’ x .583’’ surface ◦ Highly concentrated group of molecules hit only one half of the surface
Magnetic torque must be greater than this torque for optimal attitude control at this altitude
Attitude Control – Calculating Drag Force Torque
𝜏= 𝐹 𝑟= ൬𝐹𝐴 ൰𝐴𝑟 ሾwhere r is the lever armሿ
൬𝐹𝐴 ൰= .02dynescm2 = .002 Nm2
𝐴= ሺ.5∗10′′ሻሺ.583′′ሻ= 2.915 in2 = .00188 m2
𝑟= .5∗10′′ = 5 in = .127m
𝜏= ቀ𝐹𝐴 ቁ𝐴𝑟 = ቀ.002 Nm2ቁሺ.001888 m2ሻሺ.127mሻ= 4.8× 10−7 N∙m ==> 𝜏= .48 μN∙m
13
Proposed experimental setup:1) Measure the torsional spring constant of a thin stainless steel
wire
2) Hang magnet from wire and find equilibrium point at which torsional torque equals magnetic torque
3) Using the equation we will know the value of the magnetic torque.
4) Using we can find the value of the dipole moment (mu).
5) Find the magnetic torques at every location of the orbit6) Repeat process to finalize magnet choice
Attitude Control – Choosing a Magnet
𝑘 = 4𝜋2𝐼𝑑𝑖𝑠𝑘𝑇2 𝑤ℎ𝑒𝑟𝑒 𝐼𝑑𝑖𝑠𝑘 = 12𝑚𝑅2 [4]
𝜏= −𝑘𝜃
𝜏= 𝜇Ԧ × 𝐵ሬԦ
14
Experimental Refinement Oscillation Damping
◦ Viscous fluid◦ Hysteresis Rod
Magnetic Placement Orbit Simulation
◦ Time Spent / Usefulness tradeoff
Attitude Control – Issues to Consider
15
[1] http://oceanexplorer.noaa.gov/explorations/05galapagos/logs/dec22/media/magfield_600.html
[2] Bopp, Matthew, and Jonathan Messer. An Analysis of Magnetic Attitude Control of Low Earth Orbit Nano-satellites with Application for the BUSAT. BUSAT. Attitude Control and Determination Subsystem. Web. 01 Mar. 2010.
[3] Fundamentals of Space systems
[4] Physics for Scientists and Engineers (Giancoli)
Works Cited
16
DESIGN 1: • ROUGH TUBESAT DIMENSIONAL CONSTRAINTS• THIN PC BOARD, TAPERED EDGES
1.2 Mechanical Design
DESIGN 2: • SOLAR PANEL DIMENSIONAL CONSTRAINTS• TWIN HINGED BOARDS• MAXIMIZE PANEL #
DESIGN 2: • SOLAR PANEL DIMENSIONAL CONSTRAINTS• TWIN HINGED BOARDS
DESIGN 2: • SOLAR PANEL DIMENSIONAL CONSTRAINTS• TWIN HINGED BOARDS• MAXIMIZE PANEL #
DESIGN 3: • SOLAR PANEL + ANTENNAE CONSTRAINTS• CASING (E&M SHIELDING)• # PANEL REDUCTION
DESIGN 4: • REPLACE WIRE-WRAP ANTENNAE WITH PATCH• PRESERVE ENCLOSURE SYMMETRY• # PANEL REDUCTION
Purpose:◦ Fly sensors in a Low Altitude Orbit◦ Observe “Good Science” from these sensors◦ Basic Purpose of Flying a Satellite
What We Will Be Flying:◦ Two Plasma Probes◦ One Magnetometer
1.6 Sensors
22
What We Expect to Find
Current FromCharged
Particles
Graph Shows Current vs. Swept Bias Voltage
23
Plot of a Log Scale
Electron Temperature◦ Temperature of Given
Electron Distribution Plasma Potential
◦ Average Electric Potential Between Particles
What Do We Use This For?
24
Sensor System Block Diagram
To Command Interface
To Command Interface
Plasma Readings Magnetic Field Readings
Able to Connect Directly to Main Interface
Built on Individual Circuits◦ Ease of
Transferability◦ Redundant System
25
Deployable Sensor Booms◦ Fiberglass Rod◦ Langmuir Plasma Probes◦ Magnetometer
Folding Sensor Booms Plate Plasma Probes
◦ 1cm x 3cm Gold Plated Units◦ Coaxial Cable Directly Through Wall◦ Opposite Corners of Satellite
History of Design
26
Electrical Schematic
27
Past:◦ Decide on Final Design of Probes
Right Now:◦ Have Electrical Schematic◦ Have most of the parts
Next Step:◦ Assemble Circuitry◦ Test Circuitry
Timeline
28
Collecting Good Data◦ Staying Out of Wake◦ Far enough away from craft◦ Transmitting Data Back to
Earth for Analysis◦ Long Enough Orbit for Good Results◦ Stable Flight Thanks to Attitude Control
Issues/Risk Assessment
29
Power management concept◦ Energy supplied by GaAs solar panels◦ Energy stored in batteries◦ Energy provided to entire electrical system for in-
flight operation
1.5 Power Management
30
Goals:◦ Determine power supply capabilities of solar
panels and batteries◦ Regulate power usage in the satellite for
maximum data acquisition/transmission
Requirements◦ Supply sufficient power for operation of essential
satellite systems◦ Sustain power supply for estimated 3 month flight
Power Management Objectives
31
Power Supply Diagram
32
Power Usage Diagram
33
Batteries◦ 4V Batteries◦ Rated for 875mAhr
(3500mWhr) Solar Cells
◦ Rated for 14mA/ square cm at 2V
◦ Our cells can provide 400mA max (800mW)
Power Supply Specifications
34
Assumptions used to create a baseline power supply estimate:
◦ Solar panels produce full current when pointed at the sun within a 45 degree angle.
◦ Atmospheric reduction of solar energy is negligible.
◦ Satellite follows a polar orbit.
◦ Satellite attitude is primarily controlled by a fixed magnet aligning with earth’s magnetic field.
Solar Panel Power Estimation
35
Baseline Solar PowerFor a noon-midnight orbit satellite magnetic control causes solar panels to point away from the sun for a portion of a noon-midnight orbit in addition to the significant portion of orbit behind the earth’s shadow.
For a dawn-dusk orbit the sun’s rays come out of the screen and thus hit the satellite for its entire orbit.
Earth’s Shadow
Orbital Path
Sunlight
36
Baseline Solar Power II
Magnetic Field Line
Sunlight
Solar Cells Parallel to Sunlight Solar Cells Perpendicular
to Sunlight
With a single fixed magnet to control attitude, the satellite is free to rotate around magnetic field lines. Even when the field is perfectly perpendicular to the panels the rotation could cause the cells to see sunlight only 50% of the time (two sides have panels giving a 90 degree angle of effectiveness).
Direction of Satellite
37
Using all the previously discussed estimation factors, the baseline or minimum expected solar power can be calculated.
Baseline Solar Power III
Noon-Midnight Dawn-Dusk
Rotation factor 0.5 0.5
Earth's Shadow Factor 0.5 1
Magnetic Field Line Factor 0.5 1
Baseline Power Factor 0.125 0.5
Power for 5 cells (W) 4 4
Baseline Average Power (W) 0.5 2
38
Based on our Solar Power estimates, the average solar power supply should be roughly 0.5W, but this will not be continuously available.
Our batteries must store power and supply it when the solar panels are inactive.
The number of batteries launched will depend on the space and weight restrictions of our satellite after other components are installed.
Batteries
39
Power Usage SpecificationsItem Qty Voltage (V) Mode Current (mA) Duty cycle Power (mW)
Command Interface 1 5 On 10 1 50
GPS 1 5 Acquisition 29 1 145
Tracking 25 0 0
Off 0.01 0 0
MaxStream 1 5 Receive 80 0.0056 2.2
Transmit 730 0.0014 5.1
Off 0.147 0.9931 0.7
Sat Transmitter 1 3.3 Process 120 0.03 11.9
Transmit 500 0.03 49.5
Wait between Tx 0.01 0.01 0.0
Off 0.006 0.91 0.0
Magnetometer 1 5 On 35 1 175.0
Plasma Probe 2 On 1 70
Total = 510 mW
40
Our solar power estimates may prove to be too high for our real orbit.
Our transmission hardware requires large amounts of power compared to our supplies.
Aerodynamic forces may prevent rotation around the magnetic field lines resulting in solar cells never facing sunlight for up to a three month period.
Potential Issues
41
Refine Power Supply Estimate◦ Measure actual solar cell power output over
varying solar incidence angles◦ Refine orbit model following finalized attitude
control design◦ Scheduled for the first 2 weeks of April
Optimize component duty cycles Construction/Integration
◦ Install power supply systems◦ Test functionality◦ Scheduled for final 2 weeks of April, first 2 weeks
of May
Future Work
42
1.3 ESAT Communications System
Primary Link
Axonn satellite module
Frequency range: 1611.25 – 1618.75 Mhz ( Globalstar )
Data rate: 9600 144 Byte packet burst mode.
Antenna: Compact microstrip patch antenna L1 band Gain: 5.7dB 25mm x 25mm x 2mm
Current: 500mA (Tx)
Secondary Link
Maxstream spread spectrum module
Frequency range: 902 – 928Mhz ( ISM band )
Data rate: 9600 – 57.6kb
Data encryption: 32bit
Antenna: Collinear 7.5dB
RF Power: 1W
Current: 700mA (Tx)
Inventek GPS Module
Firmware: Taylor HankEYE V2.1E ( No restrictions )
Channel: 20
Update rate: 20Hz
Data rate: 115.2kb
Current: 25mA
Antenna: Active patch L1 band 28db gain 25mm x 25mm x 2mm
43
ESAT Communications System
Globalstar satellite network
ESAT
Globalstar Ground Station
Taylor Ground StationInternet
900Mhz
1611Mhz
44
ESAT Communications System
Taylor ground stationCommunications: Dual Maxstream 900Mhz ISM ModuleTracking: Az / Elv satellite antenna tracking systemAntenna: 47dB Axial mode helical stackSoftware: Sequel server database / LabView user interface / AGI satellite interface
Communications protocol HawkEYE packet structure ( high speed micro burst packet )
Packet size: 44 ByteCRC: 16 BitGPS positionInstrument dataSystem data
Byte Count 0 - 1 2 - 3 4 5 6 - 41Definition SYNC ID CMD Digital Payload Analog Payload 1 - 18Number of Byte 2 Byte 2 Byte 1 Byte 1 Byte 36 ByteRange EE EE HI Byte LO Byte HI Byte LO Byte HI Byte LO Byte HI Byte LO Byte
00 00 - FF FF 00 - FF 00 - FF 00 00 - FF FFDIP Table ID CMD = 1A DIAIP Table ID CMD = 1A DI A1 - A18
GPS Table ID CMD = 2A DI LAT, LAT REF, LON, LON REF, SPD, HDG
45
ESAT Control Module
PIC18F2620
Plasma Probe
Magnetometer
Ref.
Signal
Cal
Ramp
U/P
Clk
Temp
RX
TX
GPS Receiver
RX
TX
Extend 900
RX
TX
RX
Sleep
STX2SAT
TX
RTS
CTS
Busy
V_Solar_2
V_Solar_1
V_Batt
V_Buss
I_Buss
+5V
+5V+3.3V
+5V
+5V
Active Patch Antenna
Collinear Antenna
Patch Antenna
20Mhz Clk
46
Attitude Control◦ Magnet Finalization◦ Experimental Refinement
Mechanical◦ Drawing Finalization◦ Enclosure
Sensors◦ Circuitry Finalization
Power Management◦ Final Power Calculations◦ Circuitry Construction
Communication◦ Circuitry Design and Testing
Next Steps
47