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