1. Solar Sail Project AEM 4332W – Spacecraft Design
Preliminary Design Review
March 27, 2007
Eric Blake
Jon Braam
Raymond Haremza
Michael Hiti
Kory Jenkins
Daniel Kaseforth
Brian Miller
Alex Ordway
Casey Shockman
Lucas Veverka
Megan Williams (Team Lead)

2. Team Organization Systems Integration & Management: Megan Williams
Orbit Control: Eric Blake, Daniel Kaseforth, Lucas Veverka
Structures: Jon Braam, Kory Jenkins
Attitude Control: Brian Miller, Alex Ordway
Communications: Casey Shockman
Thermal: Raymond Haremza
Power: Michael Hiti
AEM 4332W - Solar Sail

3. Presentation Outline Project Overview Design Strategy Subgroup work
Orbit
Structure
Attitude and Control
Communication
Thermal Analysis
Power
Demonstration
Acknowledgements
AEM 4332W - Solar Sail

4. Project Overview Orbit Non-Keplarian orbit Structure
Inclination 60°
Semi-major axis 0.48AU
Structure
Target mass: 500 kg
Sail size: 100m x 100m
Inflatable boom structure, heated curing
Attitude Control
Sliding mass configuration with secondary tip thruster control
Interstellar compass – primary ADS
Communications
Ka-Band (32 GHz) Horn antennae
Thermal
Carbon mesh sail material
Multifunctional Structure (MFS) configuration
Power
Power Requirements approximately 878 W
Normal Pointing Solar Array area: meters
Silver-Cadmium (Ad-Cd) battery mass: kg
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5. Design Strategy
AEM 4332W - Solar Sail

6. Design Strategy Orbit Trade Studies Structure
Sail area versus transfer time
Varied sail size and ran simulation
Larger sail results in a faster transfer
Transfer maneuver variations
Comparison between “hot”, “cold” and simultaneous transfer trajectories
“Hot” transfer is quickest but may not be feasible due to thermal restrictions
Structure
Zero level sizing based on existing designs
Deployable space structure types
Method of rigidizing inflatable structure
Stress analysis
Determine power/time for boom deployment
Coordinate with Attitude Control and Power subgroups
Solid Modeling
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7. Design Strategy Attitude Control Trade Study Simulink modeling
Sliding mass vs Tip thruster ACS
Simulink modeling
Communication
Researched communication devices
Thermal
Solar Sail material: Mylar vs Carbon fiber mesh
Research into thermal management of spacecraft
Power
Zero level sizing for power requirements
Normal vs. Conformal Solar Array
Solar Array sizing
Battery sizing
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8. Orbit Control Eric Blake (Simulation)
Daniel Kaseforth (Control Law –Simulation )
Lucas Veverka (Control Law – Orbits)
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9. Orbit Control Problem Assumptions
How to get from Earth’s orbit to an orbit about the sun with inclination of 60° and semi-major axis of 0.48 AU using solar pressure?
Assumptions
Gravity and solar pressure are only forces
Sail is rigid flat plate and does not degrade
Sail material is perfectly reflecting
Instantaneous change in sail orientation
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10. Orbit Control Technical flow of work Simulation
Two-body force interaction (Sun, spacecraft)
Force of gravity
Force of Solar pressure
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11. Orbit Control Control Law Cone and clock angle equations
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12. Orbit Control
“Cold” orbit transfer
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13. Orbit Control
Orbital elements
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14. Orbit Control Conclusions Recommendations for further work
Simulation works
Control law functions as desired
Recommendations for further work
Sail shape analysis
Optimize transfer trajectory
Simulate sail degradation effects
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15. Orbit Control FDR Presentation
Discuss control law and simulation assumptions.
Discuss possible transfer orbits.
Show simulation results.
Justify selected transfer orbit.
Discuss further work.
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16. Structural Design
Jon Braam
Kory Jenkins
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17. Solar Sail Structure and Deployment
Challenge:
Design a deployable structure to support the sail and deliver a scientific payload.
Solution:
The sail support structure consists of four inflatable, rigidizable booms attached to a payload module.
Based on L’Garde solar sail demonstrator design.
AEM 4332W - Solar Sail

18. Aluminum Module Aluminum Unistrut Unistrut Washer Titanium Hardware
Ti Weld
Unistrut Washer
Titanium Hardware
Rubber Washer
Vibration Damping
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19. Hexagonal Shape Maximize area inside capsule
Maximize packing area inside module
Allowable surface area for features
Antenna
Camera
Solar Panel Attachment
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20. Sail Mount Hexagonal Shape Center Hole Mounting Strength Routing FEA
Add Gussets
Starburst Mount
Add connections
Center Hole
Routing
Wiring
Propellant
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21. Boom Geometry Packing constraints require tapered geometry.
Laminate thickness t = 0.25 mm.
r = 10 cm.
R = 16 cm.
l = 30 cm.
n = number of folds.
L = 72 m.
Mass ≈ 20 Kg/boom
R = r + t ( l/L)
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22. Estimate Worst Case Loading
Assumptions:
Solar Pressure at 0.48 AU = 19.8 µN/m^2.
Tip thruster forces of 150 µN.
Worst case force = 0.05 N.
Deployment load of 20 N in compression.
Thin wall tubes.
Sail quadrant loading is evenly distributed between 3 attachment points.
Quadrant area 2500 m^2.
Homogeneous material properties.
Safety factor of 3.
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23. Boom Material [0/90] carbon fiber laminate.
Polymer film inflation gas barrier.
IM7 carbon fiber, E = 276 GPa.
Low CTE.
TP407 polyurethane matrix, E = 1.3 GPa.
Tg = 55 degrees C.
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24. Expected deployment loads of 20 N in compression dictate boom sizing.
Conclusion: Booms sized to meet this requirement easily meet other criteria.
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25. Deployment Booms heated to 75 degrees C.
Inflation gas pressurizes booms for deployment.
Booms rigidize as they cool to Sub-Tg (glass transition) temperatures.
Deployment speed is controlled by a single motor which pays out the tensioning cables at 1 cm/sec.
Motor retracts tension cables after booms are rigidized to pull out the sail.
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26. Deployed Boom with Micro PPT Tip Thrusters
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27. Future Work and FDR Deliverables
Sliding mass
Size
Placement
Effects of structural deformation on attitude control.
Investigate low frequency vibration modes.
Volume of inflation gas needed.
Proper laminate analysis.
FDR Deliverables:
Configuration: Solid Model stowed and deployed
Total Mass/Moment of inertia
Deployment Methodology
Structural Analysis
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28. Attitude Control
Alex Ordway
Brian Miller
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29. Attitude Control Detailed description of trade study
Sliding Mass characteristics
Power consumption
10 W
Approximate control torques
Being calculated; will be sufficient
Mass required
10 kg, open for refinement
Tip thruster characteristics
Power Consumption
100 W
10 kg
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30. Attitude Control Detailed description of ACS
Primary use of sliding mass
Tip thrusters utilized as secondary ACS
Configuration chosen for a number of reasons
Thrusters require more power to operate (~1kw)
Ion ejection from ions could interfere with solar arrays
Operational life of thrusters limited to 3000 hours
Sliding mass offers comparable transfer times without aforementioned drawbacks
Tip thrusters chosen offer smaller force at lower power usage, no significant life restrictions, lower probability of system interference
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31. Attitude Control Detailed description ACS cont… Tip Thruster Selection
Micro Plasma Pulsed Thruster (Micro PPT)
Solid polymer fuel bar
Eliminates need for auxiliary fuel transport infrastructure
Can be utilized in off-nominal attitude situations in addition to being an available ACS when the solar sail is not deployed
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32. Attitude Control ADS Primary Interstellar Compass (ISC)
Low power
3.5 W
Exceptional Accuracy
0.1 deg (1σ)
Low mass
2.5 kg
Technology has not flown
Developed by Draper Laboratory
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33. Attitude Control ADS Secondary Sun Sensors
Located on all solar oriented exterior planes
Reorient space craft in off-nominal attitude situations
Provide data to orient solar arrays for optimal solar collection
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34. Future Work Finish attitude control simulation
Calculate final required mass for ACS
Refine simulation using information from structures group
Consider sail ejection once orbit is achieved
Independent module ACS
Reaction wheels most likely candidate
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35. Communications
Casey Shockman

36. Frequency X-Band: 8.4 GHz Ka-Band: 32 GHz
This is the typical frequency used, so DSN is becoming overloaded at this frequency.
Ka-Band: 32 GHz
Due to overloaded X-Band frequency, the DSN is migrating to Ka-Band frequency.
Can transfer data much more quickly than X-Band.
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37. Antenna Horn High data transfer rate with low power required.
Works directly with recently developed Small Deep Space Transponder.
New design works with X-Band and Ka-Band transmit as well as X-Band receive.
Lighter and smaller than parabolic reflector or array.
High gain.
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38. AEM 4332W - Solar Sail

39. AEM 4332W - Solar Sail

40. Current/Future Work
Currently, I am working on a design space to optimize values for power required, antenna sizing, pointing accuracy, and signal to noise ratio.
Problems include finding accurate equations for horn antenna systems.
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41. Thermal Analysis
Raymond Haremza
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42. Carbon Fiber Mesh Carbon Fiber Mesh developed by ESLI
Mesh is composed of a network of carbon fibers crisscross linked into a matrix that is mostly empty space.
200 times thicker than the thinnest solar sail material, but so porous that it weighs the same
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43. Common Problems Traditional materials tear easily
require heavy support structure to maintain tension
can build up static electricity
UV degrades and melt at high temperatures
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44. Carbon Fiber Mesh Can tolerate temps as high as 4,500 deg F
Small areal mass density: 30μm thickness compared to 2μm with same area density (~5g/m^2)
Immune to UV degradation
Ability to self-deploy, the carbon scrub-pad material could be packed so it pops out flat once released. This can eliminate the need for any complicated mechanical deployment mechanism, which decrease mass of the craft.
Easier to deploy because it doesn’t cling or wrinkle
Higher Melting Point
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45. Carbon Fiber vs. Traditional Material
Aluminum Coated Mylar
Using sample microtruss which is formed
from perfectly electrical conducting (PEC)
wires. The time-average force on the sail
can be found using physical optics assuming
microtruss is illuminated by a uniform plane
wave (UPW) and
Force at 0.48AU = 0.348N
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46. Thermal Analysis of Payload Module
Found an innovative way to configure spacecraft parts which eliminate chassis, cables and connectors.
MFS (Multifunctional Structures) achieves this by using MCM (multichip modules) and dissipating its heat through a thermal core fill, and utilizing aluminum honeycomb sandwiched between 2 fiber reinforced cyanate ester composite faceplates.
This high density configuration increases payload-mass fraction and provides major weight volume and cost savings.
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47. MFS Configuration Thermal copper strap used to transfer heat to
radiator surface.
Hi-K facesheets
(K13C2U)
Multichip Module -
Specialized electronic
package where multiple
integrated circuits
are packaged to do
many jobs with one
module.
High Conductivity Filler
Kz = 700 W/mK
Aluminum Honeycomb
High K Isotropic Carbon-
Carbon Doubler
Edge corefill
AEM 4332W - Solar Sail Kz

48. Thermal Control of MFS
In order to dissipate waste heat from the MCM along with solar energy loads on the outer skin.
Rate of heat flow
Effective rad environment
Radiation Equation
Emissivity of radiator
Temp of base plate
Lateral Conductance
Heat flow path length
Setting equal and solving for temp of baseplate yields
Average radiator temp
Cross sectional area
Material thermal conductivity
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49. Thermal Control Configuration Options For MFS Integration
Incorporate Thermal Doubler Hi-K Corefill Kx, Ky > W/mK; Kz> W/mK
High Conductivity
Composite
Facesheet
With Kx,Ky>
150 W/mK
Incorporate
Heat Pipes
Incorporate Deployable Radiator
Where,
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Source: Thermal Management For Multifunctional Structures

50. Confirmation of MFS
The Multifunctional Structure was successful based on the data returned from the Deep Space 1 mission. This mission the MFS was tested by powering it up once every two weeks which provided a data set containing health and status information, electrical-conductivity test data, and thermal-gradient measurements. The thermal-gradient data proved to stay within operating conditions.
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51. Thermal Analysis of Boom Supports
Carbon fiber booms need to maintain temperatures below 40 deg C.
To achieve this a coating will be applied to the outside of the carbon fiber.
By using the radiation equation and basic thermodynamics the required coefficient of absorbtivity, emmisivity can be found that satisfy these constraints. From these coefficients a coating can be chosen.
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52. Thermal Properties of Carbon Fiber Boom
Setting equal to each other and solving for
temperature of the surface of the boom yields:
Carbon fiber properties:
R=13 cm
Thickness = .25mm
Length = 72 m
K = 400 W/m
Density = 1490 kg/m^3
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53. Material Selection for Boom
By graphing different values of absortivity and emmisivity the proper
Coating can be found that will keep the boom under 313K.
White Paint S13G-LO with , gives T =282K

54. Future Work
I plan on further investigating and analyzing the spacecraft components such as the fuel tank, additional thermal control methods, and complete analysis of MFS integration into the spacecraft configuration.
Also working together with orbit group to run simulations with Aluminized Mylar, Kapton and Carbon Fiber solar sail and find best material for our mission.
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55. Power
Michael Hiti

56. Objectives
Determine the amount of power required to support the payload, and all other components of the spacecraft.
Perform a trade study to determine whether to use a normal-pointing solar array or a fixed solar array.
Determine the size and type of the solar array
Determine the size and type of the batteries that will be used
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57. Power Requirements BOL Power Requirement : ~878W
Power (W)
Remote Sensing Instruments
Coronagraph 4
All Sky Camera 5
EUV Imager 6
Magnetograph-Helioseismograph
IN-SITU Instruments
Magnetometer 2
Solar Wind Ion Composition and Electron Spectrometer
3.5
Energetic Particle 3
Communications
Satellite/Data Transmission 50
Attitude Control
125
Structure
Heat Curing Booms
675
Misc
Sliding Mass, Adjusting Array/Satellite/Antenna
TOTAL
877.5
BOL Power Requirement : ~878W
EOL Power Requirement: ~ 203W
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58. Normal Pointing Solar Array
Benefits:
A fold out array can be used to utilize its reflectance and thermal characteristics for thermal management
A sun tracker will already be being used
Able to collect maximum possible solar energy
Panels could be positioned to minimize thermal and radiation damage
AEM 4332W - Solar Sail

59. Solar Array Sizing General Formulas:
Pchg = Vchg* Ichg = (Vchg* Cchg)/15h
PEOL = PL + Pchg
PEOL = ηrad* ηangle* ηtemp* PBOL
Aarray = PBOL / (ηGaAs* IS * ηpack)
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60. Solar Array Sizing Normal Pointing Array Assuming:
a temperature efficiency reduction of ~40%
a radiation degradation of ~50%
a packing efficiency of ~90%
Gallium Arsenide cells
Approximate Solar Array Area: 2.39m^2
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61. Solar Array Sizing Conformal Solar Array Assuming:
a temperature efficiency reduction of ~55%
a radiation degradation of ~55%
cosine loss of ~81%
a packing efficiency of ~90%
Gallium Arsenide cells
Approximate Solar Array Area: 4.37m^2
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62. Cchg = (PL* td ) / (Vavg * DOD)
Battery Sizing
General Equations:
Cchg = (PL* td ) / (Vavg * DOD)
Ebat = Cchg * Vavg
mbat = Ebat / ebat
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63. Battery Sizing
Ag-Cd batteries will be used for their reasonable energy density and cycle life
Assuming:
a bus voltage of 28V
a DOD of ~25%
a maximum load duration of 2.0h
Battery Mass = kg
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64. Components Spectrolab Cells and Panels 28.3% efficiency
84 mg/cm^2 (cells)
2.06 kg/m^2 (panel)
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65. Components Moog Solar Array Drives Two-axis solar array drive
Power = 4W per axis
Mass = 4.2 kg
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66. Demonstration
For FDR, we plan to have a demonstrated orbit which includes pointing requirements and attitude control.
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67. Acknowledgements Stephanie Thomas, Princeton Satellite Systems
Professor Joseph Mueller, University of Minnesota
Professor Jeff Hammer, University of Minnesota
Dr. William Garrard, University of Minnesota
Kit Ruzicka, University of Minnesota
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