Solar Sail Project AEM 4332W – Spacecraft Design

Eric BlakeJon BraamRaymond HaremzaMichael HitiKory JenkinsDaniel KaseforthBrian MillerAlex OrdwayCasey ShockmanLucas VeverkaMegan Williams (Team Lead)

Solar Sail Project AEM 4332W – Spacecraft

Design

Preliminary Design ReviewMarch 28, 2007

AEM 4332W - Solar Sail 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


Presentation Outline

• Project Overview• Design Strategy• Subgroup work

– Orbit– Structure– Attitude and Control– Communication– Thermal Analysis– Power

• Demonstration• Acknowledgements


Project Overview

Top Level Requirements• The payload mass is 34 kg • The payload average power draw is

24.5 Watts • The final orbit should have a semi-major

axis of 0.48 AU and an inclination of 60 deg

• The launch vehicle will provide a hyperbolic escape velocity of 0.5 km/s. A Delta II 7425 will be used for launch.

• The structure will fit inside the selected launch vehicle.


Project OverviewMAJOR TASKS1) Develop control law for semi-major axis change and inclination

change to determine solar sail orientation. 2) Analyze transfer time for different sail sizes to determined

optimum sail size. 3) Conduct a trade study between sliding mass and tip thruster

attitude control systems.4) Determine the data transfer rate and power requirements for

data downlinks to Earth. Assume 2 downlinks per week to the DSN.

5) Conduct a trade study between conformal solar array and normal-pointing solar array.

6) Size the solar array to meet total power requirements.7) Analyze the thermal properties of the solar sail spacecraft.8) Choose a configuration and compute the total mass and

moment of inertia.9) Design a payload module.10) Design for the satellite actuation.11) Calculation and testing of attitude control law.


Project OverviewOrbit• Non-Keplarian orbit

– Inclination 60°– Semi-major axis 0.48AU

Structure• Target mass: 500 kg• Sail size: 100m x 100m• Inflatable boom structure, heated curingAttitude Control• Sliding mass configuration with secondary tip thruster control• Interstellar compass – primary ADS Communications• Ka-Band (32 GHz) Horn antennaeThermal• Carbon mesh sail material• Multifunctional Structure (MFS) configurationPower• Power Requirements approximately 878 W• Normal Pointing Solar Array area: 2.39 meters• Silver-Cadmium (Ad-Cd) battery mass: 13.92 kg


Design StrategyOrbit• Trade Studies

– 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• Trade Studies

– 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


Design StrategyAttitude Control• Trade Study

– Sliding mass vs Tip thruster ACS• Simulink modelingCommunication• Researched communication devicesThermal• Trade Study

– Solar Sail material: Mylar vs Carbon fiber mesh• Research into thermal management of spacecraftPower• Zero level sizing for power requirements• Trade Study

– Normal vs. Conformal Solar Array• Solar Array sizing• Battery sizing


Orbit Control

Eric Blake (Simulation)Daniel Kaseforth (Control Law –Simulation

)Lucas Veverka (Control Law – Orbits)


Orbit Control

• Problem– 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


Orbit Control• Technical flow of work

– Simulation• Two-body force interaction (Sun, spacecraft)

– Force of gravity

– Force of Solar pressure

2r

mGMF satsun

grav

sailsolar Ar

rcF

2

00


Orbit Control

– Control Law• Cone and clock angle equations

f

fefe

fefe

cos12

sincos1

4

sincos1

893

tan

2

1


Orbit Control

“Cold” orbit transfer


Orbit Control

Orbital elements


Orbit Control

• Conclusions– Simulation works– Control law functions as desired

• Recommendations for further work– Sail shape analysis– Optimize transfer trajectory– Simulate sail degradation effects


Orbit Control

• FDR Presentation– Discuss control law and simulation

assumptions.– Discuss possible transfer orbits.– Show simulation results.– Justify selected transfer orbit.– Discuss further work.


Structural Design

Jon BraamKory Jenkins


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.


Aluminum Module

• Aluminum Unistrut– Ti Weld

• Unistrut Washer• Titanium Hardware• Rubber Washer

– Vibration Damping


Hexagonal Shape• Maximize area inside

capsule

• Maximize packing area inside module

• Allowable surface area for features– Antenna– Camera– Solar Panel Attachment


Sail Mount

• Hexagonal Shape– Mounting– Strength

• FEA– Add Gussets– Starburst Mount

» Add connections

• Center Hole– Routing

• Wiring• Propellant


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)


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.


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.

mmffL EVEVE 1

m

m

f

fT E

V

E

VE mmffc VV


0 10 20 30 40 50 60 70 800

200

400bending stress

x (m)

stre

ss (

KP

a)

0 10 20 30 40 50 60 70 800

200

400buckling

x (m)

P c

ritic

al (

N)

0 10 20 30 40 50 60 70 800.5

1

1.5shear

x (m)

She

ar S

tres

s (K

Pa)

0 10 20 30 40 50 60 70 800.5

1

1.5max inflation pressure

x (m)

burs

t pr

essu

re (

MP

a)

• Expected deployment loads of 20 N in compression dictate boom sizing.• Conclusion: Booms sized to meet this requirement easily meet other criteria.


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.


Deployed Boom with Micro PPT Tip Thrusters


Future Work and FDR Deliverables

• Future Work:– 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


Attitude Control

Alex OrdwayBrian Miller


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• Mass required

– 10 kg


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


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


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


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


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

Communications

Casey Shockman


Frequency

• X-Band: 8.4 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.


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.


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.


Thermal Analysis

Raymond Haremza


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


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


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


Carbon Fiber vs. Traditional Material

Fz A0 E inc 20.8N

Using sample microtruss which is formedfrom perfectly electrical conducting (PEC) wires. The time-average force on the sailcan be found using physical optics assumingmicrotruss is illuminated by a uniform planewave (UPW) and

E inc 3kV /m

Carbon Fiber Aluminum Coated Mylar

0.00E+00

5.00E-02

1.00E-01

1.50E-01

2.00E-01

2.50E-01

3.00E-01

3.50E-01

4.00E-01

7.00E+109.00E+101.10E+111.30E+111.50E+11

Distance From Sun

Forc

e (

N)

Force at 0.48AU = 0.348N

0 8.89 10 12


CP1 Solar Sail Material

• Developed by SRS Technologies created a 5 micro meter thick film constructed of CP1 with an aluminized front surface and a black emissive black surface.

• CP1 is a unique polymer which has favorable structural characteristics.

T(K) S

D2(l b )

14

Source: Scalable Solar Sail Subsystem Design Considerations


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.


MFS Configuration

Thermal copper strap used to transfer heat to radiator surface.

Multichip Module -Specialized electronic package where multipleintegrated circuits are packaged to do many jobs with onemodule.

Hi-K facesheets(K13C2U)

Aluminum Honeycomb

Edge corefill

High Conductivity FillerKz = 700 W/mK

High K Isotropic Carbon-Carbon Doubler

Kz


Thermal Control of MFS

• In order to dissipate waste heat from the MCM along with solar energy loads on the outer skin.

Q AS(TR4 TS

4 )

Q KAx (Tbp TR )

L

T bpL

KAx

As (TR4 TS

4 ) TR

Q

Radiation Equation

Lateral Conductance

Setting equal and solving for temp of baseplate yields

K

Ax

TR

L

TS

Tbp

Rate of heat flow

Effective rad environment

Emissivity of radiator

Temp of base plate

Heat flow path length

Average radiator temp

Cross sectional area

Material thermal conductivity


Thermal Control Configuration Options For MFS Integration

High ConductivityComposite FacesheetWith Kx,Ky>150 W/mK

Incorporate Thermal Doubler Hi-K Corefill Kx, Ky > 150-1500 W/mK; Kz> 40-700 W/mK

Incorporate Heat Pipes

Incorporate Deployable Radiator

Where,

Source: Thermal Management For Multifunctional Structures


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.


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.


Thermal Properties of Carbon Fiber Boom

Ý Q rad

Carbon fiber properties:R=13 cmThickness = .25mmLength = 72 mK = 400 W/mDensity = 1490 kg/m^3

Intensity 4 *1026W

4 (.48AU)9594.12

W

m2

TS 89801SAS

1/ 4

Qsun Intensity * * A 89801

Qrad ASS (TS4 TR

4 )

Setting equal to each other and solving for temperature of the surface of the boom yields:

TS

t

R

Surface Temp vs Absortivity and Emmissivity

0

0.2

0.4

0.6

0.8

1

1.2

0 200 400 600 800

Surface Temp of Boom (K)

Em

msiv

ity a

nd

Ab

sort

ivit

y

Emmisivity

Absortivity

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 =252K

.2

.85


Surface Temp vs Absortivity and Emmissivity

0

0.2

0.4

0.6

0.8

1

1.2

0 200 400 600 800

Surface Temp of Boom (K)

Em

msiv

ity a

nd

Ab

sort

ivit

y

Emmisivity

Absortivity

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 =252K

.2

.85


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 Carbon Fiber, and CP1 solar sails and find best material for our mission.

Power

Michael Hiti


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


Power Requirements

• BOL Power Requirement : ~878W

• EOL Power Requirement: ~ 203W

Power (W)

Remote Sensing Instruments

Coronagraph 4

All Sky Camera 5

EUV Imager 6

Magnetograph-Helioseismograph 4

IN-SITU Instruments

Magnetometer 2

Solar Wind Ion Composition and Electron Spectrometer 3.5

Energetic Particle 3

CommunicationsSatellite/Data Transmission 50

Attitude Control 125

Structure Heat Curing Booms 675

MiscSliding Mass, Adjusting Array/Satellite/Antenna 50

TOTAL 877.5


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


Solar Array Sizing

General Formulas:

Pchg = Vchg* Ichg = (Vchg* Cchg)/15h

PEOL = PL + Pchg

PEOL = ηrad* ηangle* ηtemp* PBOL

Aarray = PBOL / (ηGaAs* IS * ηpack)


Solar Array Sizing

• Normal Pointing ArrayAssuming:– 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


Solar Array Sizing

• Conformal Solar ArrayAssuming:– 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


Battery Sizing

General Equations:

Cchg = (PL* td ) / (Vavg * DOD)

Ebat = Cchg * Vavg

mbat = Ebat / ebat


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 = 13.92 kg


Components

• Spectrolab Cells and Panels

– 28.3% efficiency– 84 mg/cm^2 (cells)– 2.06 kg/m^2 (panel)


Components

• Moog Solar Array Drives

– Two-axis solar array drive

– Power = 4W per axis– Mass = 4.2 kg


Future Work

• Refining sizing of battery and solar panel with more specific power requirements


Demonstration

• For FDR, we plan to have a demonstrated orbit which includes pointing requirements and attitude control.


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

Solar Sail Project AEM 4332W – Spacecraft Design

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