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Solar Sail
Department of Aerospace
Engineering and Mechanics
AEM 4332W – Spacecraft Design
Spring 2007
2
Solar Sailing:
3
Project Overview
– Motivation
– Scope
– Organization (tasks [%complete], groups,
[who?])
– Present the scope of your design work. What are
you setting out to do?
– Explain how you have organized the work. What
are the major tasks? What groups have you
organized your team into, and who is in each
group?
4
Team Members
Orbit: Eric Blake, Daniel Kaseforth, Lucas
Veverka
Structure: Jon Braam, Kory Jenkins
ADC: Brian Miller, Alex Ordway
Power, Thermal and Communication:
Raymond Haremza, Michael Hiti, Casey
Shockman
System Integration: Megan Williams
5
Design Strategy
Not yet complete. Needs:
– Describe all of the trade studies you are considering in this project
– Describe the trade study conclusions and any other design decisions
that you have already made
– Discuss the unfinished trade studies and what effect they will have
on your design
– Summarize the key properties of the mission (orbit, anticipated
lifetime, candidate launch vehicles)
– Summarize the key properties of the spacecraft (mass, dimensions,
peak and average power requirements, ADCS configuration, type of
propulsion system, list of any moving parts, other important info as
you see fit)
– Show a 3D diagram of the spacecraft (use a CAD package, ie Solid
Works or Pro-E)
6
Trade Study Results
7
Cost Estimate
Delta II Launch:
$42,000,000.00
Navigation System:
Carbon fiber booms:
$
250,000.00
Aluminum Bus:
$
1,200.00
2 stepper motors (sail deployment):
$
80,000.00
Heater:
Helium Tank:
Star Tracker:
$ 1,000,000.00
4 Step Motors (sliding masses):
$
160,000.00
Reaction Wheels:
$
600.00
Thrusters:
Antenna Horn:
Thermal Coating:
Sail material:
Solar Panels:
Total:
$43,491,800.00
Before Launch
$ 1,491,800.00
Orbit
Eric Blake
Daniel Kaseforth
Lucas Veverka
Eric Blake
Optimal Trajectory of a Solar Sail:
Derivation of Feedback Control Laws
10
Recall Orbital Mechanics
• The state of a spacecraft can be described
by a vector of 6 orbital elements.
– Semi-major axis, a
– Eccentricity, e
– Inclination, i
– Right ascension of the ascending node, Ω
– Argument of perihelion, ω
– True anomaly, f
• Equivalent to 6 Cartesian position and
velocity components.
11
Orbital Elements
12
Equations of Motion
 ^ 
v   2 r   2  r n  n


rv

2
^
r 
r
^
^

^
^
n  cos  r  sin  cos  p  sin  sin  p r
 = Sail Lightness Number
 = Gravitational Parameter
^
p
n
sun  line

sail
^
^
p r

^
r
13
Problem: Minimize Transfer
Time

 ^ 
H ( x,  , u )  r  v  2 v  r   2  r n  v  n  1
r



2
^
r 

^
^

r  3 v  3 5 (r  r )r  2 3 (r n)(v  n) n  2(r n) r 
r
r
r



^

v   r
^
p
By Inspection:
^
^
max{ n  v }  n  v
Transversality:
  ^ 2

  ^ 2


(
r

n
)
p

n


(
r

n
)
p

n
v
v
 r2

 2


 t t 0  r
 t t f
n
sun  line

sail
^
^
p r

^
r
14
Solution
• Iterative methods are needed to calculate costate boundary conditions.
• Initial guess of the co-states must be close to
the true value, otherwise the solution will not
converge.
• Difficult
• Alternative: Parameter Optimization.
– For given state boundary conditions, maximize
each element of the orbital state by an
appropriate feedback law.
15
Orbital Equations of Motion

x  g ( x,  ,  )
d
r3

sin( f   )W
df
p sin i
da
2 pr
p


Se
sin
f

T


df  (1  e 2 ) 2 
r
2
 r
de r 
r 
 S sin f  T 1   cos f  T e
df

p
p 

 r
d
d
r2 

cos i   S cos f  T 1   sin
df
df
e 
p

di r 3

cos( f   )W
df p
p
df
 2
dt
r
2
S
r

r2
cos 3 
p
1 e cos f
T 

r2
cos 2  sin  sin 
p  a(1  e 2 )

f




r
 r2 

1 
 S cos f  T 1   sin f  
p



 e 

W 

r
2
1
cos 2  sin  cos 
 = Sail Lightness Number
 = Gravitational Parameter
16
Maximizing solar force in an arbitrary direction
^
^
^
^
n  cos  r  sin  cos  p  sin  sin  p r
^
Maximize: aq     r  n  n  q 
2
~ ^
~
~ ^
~
~ ^
^
q  cos  r  sin  cos  p  sin  sin  p r
2
r 

Sail pointing for maximum
acceleration in the q direction:
^
p
sun  line
sail
~
 

n

^
^
p r

^
r
tan   
~
 3  9  8 tan 
2
~
4 tan 
17
Locally Optimal Trajectories
• Example: Use parameter optimization method to
derive feedback controller for semi-major axis
reduction.
• Equations of motion for a:
da
2 pr 2

df
 (1  e 2 ) 2
p

Se
sin
f

T


r


p
r
1 e cos f
p  a(1  e )
2
Feedback Law:
~
e sin f
tan  
1  e cos f
 


2
tan   
S
T 

r2
cos 3 

r
2
cos
 sin  sin 
2
~
 3  9  8 tan 
2
~
4 tan 
Use this procedure for all orbital elements
18
Method of patched local steering laws (LSL’s)
• Initial Conditions: Earth Orbit
a
1 
e
0 
 
 
i
0 

 
 
 
0 
 
0 
 
 
   t t0 0
• Final Conditions: semi-major axis: 0.48 AU
inclination of 60 degrees
a
0.48 AU 
e
 ~0 
 



i
 60 
  


free
 


 
 free 
 


   t tf  free 
19
Trajectory of SPI using LSL’s
Time (years)
20
21
Global Optimal Solution
– Although the method of patched LSL’s is not ideal, it is a solution that is
close to the optimal solution.
– Example: SPI Comparison of LSL’s and Optimal control.
22
Conclusion
• Continuous thrust problems are common in
spacecraft trajectory planning.
• True global optimal solutions are difficult to
calculate.
• Local steering laws can be used effectively to
provide a transfer time near that of the global
solution.
Lucas Veverka
•Temperature
•Orbit Implementation
Optimal Trajectory of a Solar
Sail: Orbit determination and
Material properties.
Lucas Veverka
25
Reflectivity Approximation
• Reflectivity constant, r, negatively affects the
solar radiation pressure force.
f  2PArui  n n
2
–
–
–
–
P is the solar pressure as a function of distance.
A is the sail area being struck by the solar radiation.
ui is the incident vector.
n is the vector normal to the sail.
• Emissivity and specular reflection neglected.
• Assumed a Lambertian surface.
26
Sail Surface Temperature
 Fsolar
Tsurface  
2
  4d sun





1
4
• Fsolar is the solar flux.
•
•
•
•
α is the absorptance.
ε is the emittance.
σ is the Stefan-Boltzman constant.
dsun is the distance from the sun.
27
Transfer Orbits
• Objective:
-Reach an orbit with semi-major axis of 0.48 AU
and inclination of 60 degrees as quickly as possible.
• Investigated four possible orbits
-Cold transfer orbit
-Hot transfer orbit
-Inclination first transfer orbit
-Simultaneous orbit
28
Cold Transfer Orbit
• Advantages:
– Very simple two-stage transfer.
– Goes no closer to sun than necessary to avoid
radiation damage.
• Disadvantages:
– Is not the quickest orbit available.
• Order of operations:
– Changes semi-major axis to 0.48 AU.
– Cranks inclination to 60 degrees.
• Time taken:
– 10.1 years.
29
Cold Transfer Orbit
30
Hot Transfer Orbit
• Advantages:
– Still simple with three-stages.
– Is a much quicker transfer.
• Disadvantages:
– Radiation is very intense at 0.3 AU.
• Order of operations:
– Changes semi-major axis to 0.3 AU.
– Cranks inclination to 60 degrees.
– Changes semi-major axis to 0.48 AU.
• Time taken:
– 7.45 years.
31
Hot Transfer Orbit
32
Inclination First Transfer Orbit
• Advantages:
– Very simple two-stage transfer.
– Avoids as much radiation damage as possible.
• Disadvantages:
– Takes an extremely long time.
• Order of operations:
– Cranks inclination to 60 degrees.
– Changes semi-major axis to 0.48 AU.
• Time taken:
– 20.15 years.
33
Inclination First Transfer Orbit
34
Conclusion
• Simultaneous transfer is too complicated with
little or no real benefit.
• Inclination first transfer takes too long.
• Hot transfer orbit is much quicker but submits
materials to too much radiation.
• Cold transfer orbit is slower than the hot but
gets the equipment to the desired location
safely.
• Choice: Cold transfer orbit!
Daniel Kaseforth
Control Law Inputs and Navigation
System
36
Structure
Jon T Braam
Kory Jenkins
Jon T. Braam
Structures Group:
• Primary Structural Materials
• Design Layout
• 3-D Model
• Graphics
-72 Project Hours-
39
3-D Model
• Blah blah blah (make something up)
40
Graphics
• Kick ass picture
41
Launch Vehicle
Weight and Volume Constraints
• Delta II : 7400 Series
• Launch into GEO
– 3.0 m Ferring
» Maximum payload mass: 1073 kg
» Maximum payload volume: 22.65 m3
– 2.9 m Ferring
» Maximum payload mass: 1110 kg
» Maximum payload volume: 16.14 m3
42
Launch Vehicle
2.9 Meter Ferring
In Flight Organization
– Antenna Stowed
– Solar panels folded to
sides
43
Primary Structural Material
Aluminum Alloy Unistrut®
– 7075-T6 Aluminum Alloy
• Density
– 2700 kg/m3
– 168.55 lb/ft3
• Melting Point
– 477 to 635oC
http://www.matweb.com/SpecificMaterial.asp?bassnum=MA7075T6
44
Primary Structural Material
• Density
• Mechanical Properties
– Allowing Unistrut design
• Decreased volume
• Factor of safety 15.0
– 0.06in thick (12 Ga)
• Thermal Properties
– Capable of taking thermal loads
45
Primary Structural Material
• 7075-T6 Aluminum Alloy
– Used in ISS structure
– Useful Characteristics
• Resistance to general corrosion
• Resistance to pitting
• Crack resistance
– Inter-grainulairy
– Stress corrosion
46
Design Layout
• Constraints
– Volume
– Service task
– Thermal
consideration
– Magnetic
consideration
– Vibration
– G-loading (7.5)
47
Design Layout
• Unistrut Design
– Allowing all inside surfaces to be bonded to
• Titanium hardware
– Organization
• Allowing all the pointing requirements to be met with
minimal attitude adjustment
48
Design Layout
49
3-D Model
• Sail Attachment
– Single Tube
• 0.09m Diameter
– Allows Nitrogen to feel
stabilizing thrusters
– Supports Sail and Argon
External Tank
– Mounted to tube bus
50
Odds and Ends
• Things to be improved upon
– Safety Factor of 14 in Compression and Bending
• Could be reduced to save weight
Other Projects
•Sliding mass movement design
•Boom deployment methods
•Moment of Inertia determination
•Mass budget
51
Other Projects
Trade Studies
- Structural Materials used in ISS and other
long term spacecraft.
- Deployment methods and other autonomous
movement used in space.
- In space structural connections
Kory Jenkins
• Sail Support Structure
• Anticipated Loading
•Stress Analysis
• Materials
•Sail Deployment
53
Sail Sizing
• Characteristic acceleration is a measure of
sail performance.
2P
ao 
 s  mp / A
 s  ms / A
• Characteristic acceleration increased with
sail size.
• Higher acceleration results in shorter
transfer time.
• Sail size is limited by launch vehicle size
and deployment power requirements.
54
Sail Support Structure
• Challenge: Design a robust, easy to deploy structure that
will maintain sail shape.
• A 150 x 150 meter sail covers the same area as 5 football
fields. (22,500 square meters)
• Solution: An inflatable boom structure based on the L’Garde
design supports 4 triangular sail quadrants.
• Booms are deployed in pairs to minimize power
consumption.
55
Step 5
Step 1
Deployment cables retract
to pull the sail quadrants
out of their storage
compartments.
Heater: Raises boom
temperature above glass
transition temperature to
75 C.
To sail quadrant
Step 4
Once deployed, booms cool
below glass transition
temperature and rigidize.
Step 2
Inflation gas inlet:
booms are inflated to 120
KPa for deployment.
Step 3
Cables attached to stepper
motors maintain deployment
rate of ~ 3 cm/s.
To deployment motor
56
Estimate Worst Case Loading
Solar Pressure
P = 2/3 P_quadrant
Assumptions:
• Solar Pressure at 0.48 AU
= 19.8 µN/m^2.
• Thin wall tube.
• Sail quadrant loading is
evenly distributed between
3 attachment points.
• Isotropic material
properties.
• Safety factor of 3.
57
Analysis of a Tapered Beam
My

Bending
I
2

EI
Buckling
Pcr 
4L2
Shear
VQ
 max 
Iy
Hoop stress
(inflation pressure)
Section
Modulus
Pmax 
t 
 hoopt
r
x
S ( x)  dA  (dB  dA) 
4
L
2
58
•
•
•
Expected deployment loads of 20 N in compression dictate boom sizing.
Booms sized to meet this requirement easily meet other criteria.
Verified using laminate code that accounts for anisotropy of composite materials.
59
Boom Specifications
•
•
•
•
•
Cross-ply carbon fiber laminate.
IM7 carbon fiber
TP407 polyurethane matrix, Tg = 55 deg C
Major Radius = 18 cm, minor radius = 10 cm.
Length = 106 meters.
Analysis of a Composite Laminate:
EL  V f E f  Vm Em
 V f Vm 

ET  

E

 f Em 
1
  [Q]K  [ o  z   T ]
K
60
Conclusions and Future Work
• Sail support structure can be reliably deployed and
is adequately designed for all anticipated loading
conditions.
• Future Work
– Reduce deployment power requirement.
– Reduce weight of support structure.
– Determine optimal sail tension.
Attitude Determination and
Control
Brian Miller
Alex Ordway
Alex Ordway
60 hours worked
Attitude Control Subsystem
Component Selection and
Analysis
63
Design Drivers
•
•
•
•
•
Meeting mission pointing requirements
Meet power requirements
Meet mass requirements
Cost
Miscellaneous Factors
64
Trade Study
• Sliding Mass vs. Tip Thruster Configuration
– Idea behind sliding mass
65
Trade Study
• Sliding mass ACS offers
– Low power consumption (24 W)
– Reasonable mass (40 kg)
– Low complexity
– Limitations
• Unknown torque provided until calculations are made
• No roll capability
• Initially decided to use combination of sliding
mass and tip thrusters
66
ADCS System Overview
• ADS
– Goodrich HD1003 Star Tracker primary
– Bradford Aerospace Sun Sensor secondary
• ACS
– Four 10 kg sliding masses primary
• Driven by four Empire Magnetics CYVX-U21 motors
– Three Honeywell HR14 reaction wheels
secondary
– Six Bradford Aero micro thrusters secondary
• Dissipate residual momentum after sail release
67
ADS
• Primary
– Decision to use star tracker
• Accuracy
• Do not need slew rate afforded by other systems
– Goodrich HD1003 star tracker
•
•
•
•
•
2 arc-sec pitch/yaw accuracy
3.85 kg
10 W power draw
-30°C - + 65 °C operational temp. range
$1M
– Not Chosen: Terma Space HE-5AS star tracker
68
ADS
• Secondary
– Two Bradford Aerospace sun sensors
•
•
•
•
•
Backup system; performance not as crucial
Sensor located on opposite sides of craft
0.365 kg each
0.2 W each
-80°C - +90°C
69
ACS
• Sliding mass system
– Why four masses?
– Four Empire Magnetics CYVX-U21 Step Motors
•
•
•
•
•
•
Cryo/space rated
1.5 kg each
28 W power draw each
200 °C
$55 K each
42.4 N-cm torque
70
ACS
• Gear matching- load inertia decreases by the
gear ratio squared. Show that this system
does not need to be geared.
1
2
70m  a(600sec)
2
a  0.00389 sm2
F  ma  (10kg )(0.00389 sm2 )
F  0.0389 N
71
ACS
• Three Honeywell HR14 reaction wheels
– Mission application
– Specifications
•
•
•
•
•
•
7.5 kg each
66 W power draw each (at full speed)
-30ºC - +70ºC
0.2 N-m torque
$200K each
Not selected
– Honeywell HR04
– Bradford Aerospace W18
72
ACS
• Six Bradford micro thrusters
– 0.4 kg each
– 4.5 W power draw each
– -30ºC - + 60ºC
– 2000 N thrust
– Supplied through N2 tank
73
Attitude Control
• Conclusion
– Robust ADCS
• Meets and exceeds mission requirements
• Marriage of simplicity and effectiveness
• Redundancies against the unexpected
Brian Miller
•Tip Thrusters vs. Slidnig Mass
•Attitude Control Simulation
75
Attitude Control
• Conducted trade between tip thrusters and
sliding mass as primary ACS
• Considerations
– Power required
– Torque produced
– Weight
– Misc. Factors
76
Attitude Control
• Tip Thrusters (spt-50)
– Pros
• High Torque Produced ~ 1.83 N-m
• Low weight ~ 0.8 kg/thruster
– Cons
• Large Power Requirement ~ 310 Watts
• Lifetime of 2000 hrs
• Requires a fuel, either a solid or gas
77
Attitude Control
• Attitude Control System Characteristics
– Rotational Rate
– Transfer Time
– Required Torque
– Accuracy
– Disturbance compensation
78
Attitude Control
• Requirements
– Orbit
• Make rotation rate as fast as possible
• Roll spacecraft as inclination changes
– Communications
– Within Maximum Torque
• Pitch and Yaw Axis
~ 0.34 N-m
• Roll Axis
~ 0.2 N-m
U 
mFz m – sliding mass
M F – solar force
z – distance from cg
M – spacecraft mass
79
Attitude Control
• Pitch and Yaw Axis
•
Rotation Rate = 0.144 rad/hr
~ 8.25 deg.
•
Transfer Time = 5300s ~ 1.47 hrs
•
Required Torque = 0.32 N-m
~ 98.8% of maximum produced
•
Converges to desired angle
Torque Req.
Transfer Time
Slope = 0.00004 rad/s
80
Attitude Control
• Roll Axis
Torque Req.
•
Rotation Rate = 0.072 rad/hr
~ 4.12 deg
•
Transfer Time = 7000s ~ 1.94 hrs
•
Required Torque = 0.15 N-m
~ 75% of maximum produced
•
Converges to desired angle
Transfer Time
Slope = 0.00002 rad/s
Power, Thermal and
Communications
Raymond Haremza
Michael Hiti
Casey Shockman
Raymond Haremza
Thermal Analysis
•Solar Intensity and Thermal
Environment
•Film material
•Thermal Properties of Spacecraft Parts
•Analysis of Payload Module
•Future Work
Thermal Analysis and Design
-Raymond Haremza
84
Design Approach Strategy
85
Decision to take “cold” orbit
By taking longer to get to 0.48 AU, we in
turn reduce the amount of design, analysis,
production time and weight.
Solar Sail Material and Thermal
Analysis
86
87
Payload Panel Analysis
The Carbon-Carbon Radiator has aluminum
honeycomb sandwiched between it, and
has thermal characteristics, Ky=
Kx=230W/mK, and through the thickness
Kz = 30W/mK which allows the craft to
spread its heat to the cold side of the
spacecraft, but also keeping the heat flux to
the electric parts to a minimum.
Material Properties
  0.06
  0.78
E  1.2e7 psi
G  6.11e6 psi
v  0.32
88
Spacecraft Heat Transfer Analysis
4 1026 W
flux 
 2
2
4  d
m
7.00E+03
6.00E+03
5.00E+03
4.00E+03
3.00E+03
2.00E+03
1.00E+03
0.00E+00
9.80E-01 8.80E-01 7.80E-01 6.80E-01 5.80E-01 4.80E-01
Qsun  flux  A   Watts
Distance from Sun (AU)
 Qsun 

Tsurface  
 Atotal 
1
4
 Kelvin
Solar Intensity (flux) (W/m^2)
Solar Intensity vs Distance
89
Heat Transfer Analysis
Qsun  flux   A
4
Qrad      Atot  Tsurf
Tsurf
 Qsun 

 
 Qrad 
1
4
Setting the heat fluxes together yields the
surface temperature of the object based on
emmissivity, absorbitivity, size and geometry of
the object.
Atot
A
Thermal Analysis of Payload
Module
90
Thermal Analysis of Payload
Module
91
92
Temperature vs Distance (Side of Payload
Module)
300
280
Temperature (K)
260
85
80
75
70
65
60
55
240
220
200
180
160
140
120
100
4.80E-01
5.80E-01
6.80E-01
7.80E-01
8.80E-01
Distance from Sun (AU)
9.80E-01
deg
deg
deg
deg
deg
deg
deg
93
Temperature vs Distance (Top of Payload Module)
450
400
Temperature (K)
350
0 incidence
5 deg
10 deg
15 deg
20 deg
25 deg
30 deg
35 deg
300
250
200
150
4.80E-01
5.80E-01
6.80E-01
7.80E-01
Distance from Sun (AU)
8.80E-01
9.80E-01
Spacecraft Component Thermal
Management
Notes: By using thermodynamics the amount of heat needed to be
dissipated from the component taking into account its heat generation,
shape, size, etcetera. If the component is found to be within its operating
range, the analysis is done, if not a new thermal control must be added or
changed.
94
95
Thermal Analysis of Antenna
96
Antennae Operating Temp (-373 to 373K) vs
Distance With White Paint Reflector
390
Temperature (K)
370
350
330
310
290
270
250
4.80E-01
5.80E-01
6.80E-01
7.80E-01
Distance From Sun (AU)
8.80E-01
9.80E-01
97
Star Tracker Thermal Analysis
Using the heat generated (10W), and using common coating
material ( ); the required to maintain the star tracker’s temperature
to 30 K can be found by.
Qdiss  Qtot  T Atotal
4
s
Knowing the heat needed to dissipate, a radiator size can
be calculated, or other thermal control methods (MLI) can
be used to maintain temperature.

Arad
Qsun  Qgenerated

 Atotal
4
Ts 
98
Heat Needed to Radiate Away From Star Tracker to Keep Temp
303K
1.60E+03
1.40E+03
1.20E+03
Heat (W)
1.00E+03
8.00E+02
6.00E+02
4.00E+02
2.00E+02
0.00E+00
4.80E-01
-2.00E+02
5.80E-01
6.80E-01
7.80E-01
Distance (AU)
8.80E-01
9.80E-01
99
Using the amount of heat needed to be radiated from star tracker, the
additional area required to dissipate heat can be calculated and
chosen.
Area of Radiator Needed to Keep Star Tracker Surface Temp at
303K
Area of Radiator (m^2)
2.50E+00
2.00E+00
1.50E+00
1.00E+00
5.00E-01
0.00E+00
4.50E-01
5.00E-01
5.50E-01
6.00E-01
Distance (AU)
6.50E-01
7.00E-01
100
Thermal Analysis of Microthruster
Notes: Since Microthrusters need to be within
247 to 333 K, will have to add MLI to stay
within thermal constraints.
Analysis of Multilayer insulation…
101
Microthruster and Sun Senser
Temperature vs Distance
Temperature (K)
700
650
600
550
500
Microthruster Side
Microthruster Top
Sun Sensor
450
400
350
300
250
200
4.80E01
5.80E01
6.80E01
7.80E01
8.80E01
Distance (AU)
9.80E01
102
Thermal Analysis of Solar Panels
Need to radiate heat away from solar sail, any
ideas, stephanie, group?
103
Tempurature (K)
Solar Panel Temp (Operating temp 123
to 400K) vs Distance from Sun
580
560
540
520
500
480
460
440
420
400
380
360
340
320
300
4.80E-01
5.80E-01
6.80E-01
7.80E-01
8.80E-01
Distance from sun (AU)
9.80E-01
104
Casey Shockman
• Communications
105
Major Tasks
• Trade Studies
– Frequency
– Antenna types
– Power
– Data transfer rates
• Sizing the Antennas
• Determine placement of antennas
106
Antenna Selection and Sizing
• Initial Conditions
– Payload stores data at a rate of 15.6 kbps.
– Need to transmit data 1 or 2 times per week.
• 1 week of storage is equal to around 9,500,000 kb.
• We choose two 12,000,000 kb hard drives to store
information. One hard drive will be used as backup.
– Satellite needs to transmit data anywhere from .5 to
1.5 AU
– All aspects of the DSN (size, SNR, noise temp.etc.)
107
Frequency
• S-Band: 2 GHz
– Used primarily for short distance.
• X-Band: 8.4-8.5 GHz
– This is the typical frequency used, so DSN is
becoming overloaded at this frequency.
• Ka-Band: 31.8-32.3 GHz
– Due to overloaded X-Band frequency, the DSN is
migrating to Ka-Band frequency.
– Can transfer data much more quickly than X-Band.
Solar Sail will use Ka-Band transmit with X-Band
receive/transmit capabilities.
108
Process
• This equation was then used with the
following BER vs. SNR to solve for
variables.
109
Bit Error Rate vs SNR
110
Process
• A SNR is chosen to correspond to a BER of 10-6.
• T is noise temperature which is based on the
angle with the sun and earth, elevation angle of
the earth antenna, weather conditions, distance
between satellites.
• From this, the gain and power transmitted was
optimized for each frequency, antenna, distance
and data transfer rate
• The following chart was created for each
antenna, frequency, and distance from the sun.
Variables included power, noise temperature,
and antenna size.
111
112
Antenna Types
• Directional
–
–
–
–
Parabolic Reflector
Horn
Array
Helix
• Omni-directional
– Dipole
– Conical
113
High Gain Directional Antennas
114
Directional Antennas
• Parabolic Reflector
– High data transfer rate with low power required.
– Works with either X-Band receive/transit or KaBand receive/transit, not both.
– Conventionally heavier than horn, but recent
unused membrane dish antennas may be lighter
in the future.
– Can achieve high gain and a range of
beamwidths.
115
Directional Antennas
• Arrays
– Gain is low for small areas.
– Heavier than horn or parabolic reflector due to
the large area needed to achieve desired level of
gain.
– Can attain any beamwidth.
116
Directional Antennas
• Helix
– Can attain any beamwidth necessary.
– Antenna will have a low diameter but needs to
be long to achieve high gain.
– Length of antenna makes pointing and storage
very difficult.
– Length of antenna also adds resistance, so
efficiency drops with length.
117
Directional Antennas
• 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.
– Smaller than conventional parabolic reflector and array.
– High gain.
– Ability to track using Delta Differenced One-Way Range
(DDOR) because two tones can be sent at once (DSN
stats.pdf 9).
– Small beamwidth, suitable for long-range
communications.
– The Solar Sail will have two horn antennas.
118
119
120
Conclusions
• The horn antenna was chosen because of its
small size compared to the other choices.
• The antenna cannot transmit at a Sun-EarthProbe angle smaller than .3 degrees or on a
very stormy day at the ground station.
• Different antennas would be used on the sun
side and shade side of the antenna.
• The sun side antenna would be .2 meters in
diameter. The shade side antenna would be
.075 meters.
121
More conclusions
• The minimum transfer time for this setup is 1
hour using Ka-band transmission.
• If the required signal to noise ratio is not met
due to SEP angle or weather on earth, the
transfer rate can be slowed to allow for more
accurate data.
• Power used for transfer is 30 watts.
122
Directivity
• Horn directivity is estimated by the following
equation:
225
HPBW 
 *d
123
Beamwidths
• Using this equation:
– Sun-side antenna
• X-Band HPBW=13.42
• Ka-Band HPBW=3.35.
– Shade side antenna
• X-Band HPBW=35.79
• Ka-Band HPBW=8.95
• These beamwidths are all much larger
than the pointing accuracy so there
will be very little pointing error.
124
Low Gain Omni-Directional Antennas
125
Low-Gain Antenna Selection
• Omni-Directional Antenna
– The goal is have a low data rate
communications when not pointing at earth
– There are many choices for low gain
antennas. The solar sail will have two conical
equiangular spiral antennas.
– These two antenna will ensure the satellite
will always be within contact with the DSN.
Omni-directional Transfer
Dsn stats 5
126
127
2-Arm Conical Equiangular Spiral Antenna
Gain will be 0 dBi (isotropic)
from -70 to +70 degrees.
Gain will be -25 dBi from -90 to
-70 and 70 to 90 degrees for
each antenna.
Using this configuration, at the
worst case scenario, the low
gain antenna can transmit 1
bps with an accuracy of 10-3.
128
Costs
129
DSN Cost
Dsnstats.pdf
This gives a cost of about $1100 per hour of
transmission within the DSN network.
130
Antenna Costs
• .2 m diameter horn antenna:
• .075 m diameter antenna:
• conical equiangular antenna:
• hard drive:
• Total cost:
131
Masses
• .2 m diameter horn antenna:
– 2.75 kg
• .075 m diameter antenna:
– .40 kg
• conical equiangular antenna:
– 2 x .25 kg
• hard drive:
– 2 x .79 kg
• Miscellaneous
– 1 kg
• Total mass = 6.23 kg
Michael Hiti
Power
133
Objectives
• Determine the amount of power required to support the
payload instruments, and all other components of the
spacecraft
• Perform a trade study to determine whether to use a
normal-pointing or conformal solar array
• Determine appropriate solar array materials
• Determine appropriate solar array size
134
Objectives (continued)
• Determine appropriate battery type to be used in mission
• Determine appropriate battery size
135
Power Requirements
Peak Power (W)
Remote Sensing Instruments
Coronograph
4
All Sky Camera
3
EUV Imager
5
Magnetograph - Helioseismograph
5
Magnetometer
2
IN-SITU Instrument Package
Solar Wind Ion Composition and
Electron Spectrometer
Energetic Particle (20keV - 2MeV)
3.5
2
Attitude Control
Small Reaction Wheels
70
Large Reaction Wheel
70
Sliding Mass
40
Structure
Heat Curing Elements
335
Communications
Antenna Gimbal
8
Antenna
36
Thermal Management
50
Misc/Thermal
TOTAL
633.5
• All power
requirements for
solar sail
136
Power Requirements (continued)
Peak Power
(W)
Structure
Heat Curing Elements
335
Communications
Antenna
36
Large Reaction Wheel
70
Thermal Management
50
TOTAL
491
Attitude Control
Misc/Thermal
• Anticipated beginningof-life (BOL) power
load
137
Power Requirements (continued)
Remote Sensing Instruments
Coronograph
4
All Sky Camera
3
EUV Imager
5
Magnetograph - Helioseismograph
5
Magnetometer
2
IN-SITU Instrument Package
Solar Wind Ion Composition and
Electron Spectrometer
Energetic Particle (20keV - 2MeV)
3.5
2
Attitude Control
Small Reaction Wheels
70
Communications
Antenna Gimbal
8
Antenna
36
Thermal Management
50
Misc/Thermal
TOTAL
188.5
• Anticipated endof-life (EOL)
power load
138
Array Sizing
• Key Equations
Vchg = (1.2) * Vbus= 34.2 V
Cchg = (PL* td ) / (Vbus* DOD) = 52.9 Ah
Pchg = (Vchg* Cchg)/15h = 120.6 W
PEOL = (PL + Pchg) = 310 W
•
•
•
•
•
•
Vchg is the array voltage
Cchg is the total charge capacity of the battery
PL is the required power load at EOL
td is the anticipated max load duration (2h)
Pchg is the power required to charge the batteries
DOD is the depth of discharge (0.25)
139
Array Sizing (continued)
• The BOL power requirement is found by assessing the
various efficiency factors that lead to the conditions at EOL
Temperature efficiency = ηtemp = 1 - (0.005/K)*(Tmax – Tnom)
Radiation efficiency = ηrad = 1- R
Cosine loss = ηangle = cos(α)
PEOL = ηtemp * ηrad * ηangle * PBOL
•
•
•
•
Tmax is the maximum solar cell operating temperature
Tnom is the nominal solar cell operating temperature
R is the percent loss due to radiaiation damage
α is the maximum angle off-normal to the sun
140
Array Sizing (continued)
• Using a conformal solar array
Assuming:
ηtemp ≈ 0.51
ηrad ≈ 0.3
ηangle ≈ 0.81
PBOL = 1395 W
141
Array Sizing (continued)
• Array area equations
Acell = PBOL / ( ηGaAr* Is )
Aarray = Acell / ηpack
•
•
•
•
•
Acell is the area of the solar cells
Aarray is the area of the array
ηGaAr is the efficiancy of the solar cells
Ηpack is the packing efficiency
Is is the solar intensity
142
Array Sizing (continued)
Acell = 0.8718 m^2
With a packing efficiency of 90%
Aarray = 0.969 m^2
• These values reflect the sizes required to meet EOL power requirements
at 0.48AU
• We must check to make sure this array area will generate enough power
to support the BOL requirements at 1AU
143
Array Sizing (continued)
• Assuming that there is no radiation and cosine loss
• Assuming a ηtemp ≈ 0.90
• Is = 1355W/m^2 at 1AUl
The BOL load ≈ 546W
This would require an Acell ≈ 1.413 m^2 and an Aarray ≈ 1.57 m^2
This means that the array sizing based on the EOL requirements will not
support the BOL load requirments.
• The BOL load requirements are the driving force behind the array
sizing
144
Array Mass
• Gallium Arsenide cells weigh 84mg/cm
• Solar panels and coverslides weigh 2.06 kg/m^2
• Aluminum honeycomb panel backing weighs 0.9 kg/m^2
The total mass of a conformal array will be 5.963 kg
145
Solar Array
• Solar cells and panels
made by Spectrolab
– Ultra Triple Junction GaAs
cells
– 28.5% efficiency
– 84 mg/cm^2 (cells)
– 2.06 kg/m^2 (panel)
146
Trade Study
• Advantages to using of a normal-point solar array
– Able to collect maximum possible solar energy
– Requires smaller solar array
– Array could be positioned to minimize thermal and radiation damage
• Disadvantages to using of a normal-point solar array
– Added mass of gimbal used for positional array
– Added complexity to design
– Creates problems regarding stowage in capsule
147
Trade Study (continued)
• The BOL power requirements have caused our solar array to be nearly
twice area required to meet the EOL power requirements
• The reduction of mass is our highest priority
• The smallest gimbal used for array positioning alone weighs
approximately 5kg
– This is nearly equal to the entire mass of our array
• Since our array is already oversized for EOL requirements, an array
with normal pointing capabilities will not be beneficial
148
Battery Sizing
• Key Equations
Cchg = (PL* td ) / (Vbus* DOD) = 52.9 Ah
Ebat = (Vbus* Cchg) = 1508 W h
mbat = Ebat / ebat
• Ebat is the battery energy capacity
• ebat is the energy density of the battery
• mbat is the mass of the battery
mbat = 8.6 kg
149
Battery
• Batteries made by BST
Systems
– Silver-Zinc Battery
– 1.5 V/cell
– 175(W h) / kg
150
Demonstration of Success
151
Failure Modes and Effects Analysis
•
Boom fails to fully inflate due to problem with tank, heater, etc.
–
–
Sail may still function, would apply different torques, difficult to control
One or more of the booms could fail to extend fully. i.e. the heaters don't
work, or the inflation gas tank ruptures or it gets caught on something.
If that were to happen, it might be possible to run up the sail part way,
although there would be a lot of slack in it, and therefore a loss of
propulsion efficiency. And the attitude control system might not be able to
compensate for the asymmetric torque...assuming the sliding mass on the
malfunctioning boom worked at all...I mean, um...yeah, it'll work
perfectly...
•
•
Failure of navigation system... sail fails to know it's location and can no longer implement control laws; will not reach desired orbit.
Failure Modes and Efffects:
1. Module structure fails at 7.5 g's and breaks it shit off on exit.
Effect: It will spread debris throughout LEO. Something like
the Chinese did about 9 months ago. Oops. My Bad.
2. The Sail gets kinked inside the Bus module and is unable to deploy
or rips on deployment.
Effect: Huge embarrassing failure for the UofM design team.
3. The solar array is not able to pivot downward from its
storage/capsule setup to its working format.
Effect: Same as #2.
•
FMEA
Thermal can screw everything up. I don’t think I can narrow it down to one thing. If I have to I guess I will. Anyways, heres my FDR slides thus far,
not done yet, but pretty much done calculating stuff. Now I have to explain things, add equations and graphics and explain what I would do if I had
more time. I think stephanie will have plenty to say about what I have already. Thanks
152
Future Work
153
Acknowledgements
•
•
•
•
•
•
Stephanie Thomas
Professor Joseph Mueller
Professor Jeff Hammer
Dr. Williams Garrard
Kit Ru….
?? Who else??