Formation Flying
Rachel Winters
Shunsuke Hirayama
Matt Whitten
Tsutomu Hasegawa
Kyle Tholen
Aziatun Burhan
Matt Mueller
Masao Shimada
Shelby Sullivan
Tomo Sugano
Eric Weber
1
Columbia Disaster
• Structural damage to tiles
on the inboard leading
edge of left wing,
sustained during launch
• Columbia went on to
disintegrate in re-entry
• If there had been satellite
surveillance of exterior of
shuttle, this damage would
have likely been noticed
and dealt with
Rachel Winters (2/32)
http://www.spaceref.com/news/viewnews.html?id=733
2
Damage Criteria
• Damaged tiles require attention if they are:
 0.25 inch on the wing leading edge lower surface (RED)
 0.5 inch threshold for nose cap (GREEN)
 1 inch major dimension around TPS seals and the wing leading
edge upper surface (BLUE)
 3 inch threshold for all other TPS surfaces (YELLOW)
Rachel Winters (3/32)
http://spaceflight.nasa.gov/shuttle/rtf/hill/inspect_repair_sep9_summary_covey.ppt
3
Design
• A satellite that will fly as an
escort to the space shuttle
• Satellite provides visual
inspection of shuttle
exterior for 24 hour period
of time
• Satellite will be transported
into space on shuttle
• Satellite must meet
University Nanosat
requirements for mass and
volume
x
y
Rachel Winters (4/32)
z
4
Previous Work
• AERCam “Sprint” (NASA)
http://spaceflight.nasa.gov/station
/assembly/sprint/index.html
– Successfully tested on STS-87 for 1.25 hours
around the Columbia orbiter in 1997.
– Live video feed
– Remote controlled
– 14 inches, 35 pounds
• Mini AERCam (NASA)
http://aercam.jsc.nasa.gov/
Rachel Winters (5/32)
– Successful ground tests, still under development
– Live video feed, including orthogonal view
– Remote and supervised autonomous control
options
– 7.5 inches, 10 pounds
5
Improvements
Our design is...
– Powered sufficient to operate for 24 hours
– Autonomous control
– Some supervision is advised to ensure there is no
risk of collision with the orbiter
– Supervision is only necessary for launch and
retrieval
Rachel Winters (6/32)
6
Systems Integration & Management
Rachel Winters, Matt Whitten
Major Tasks:
• Expendable vs Recoverable spacecraft
• Recovery method design
• Determine shuttle interface requirements
• Determine picture order
Rachel Winters (7/32)
7
Relative Orbit Control & Navigation
Kyle Tholen, Matt Mueller
Major Tasks:
• Determine relative orbit to meet mission
requirements
• Determine major disturbances from orbit and
counteract them
• Single vs Multiple spacecraft trade study
• Determine thruster equipment
• Find Tank size
• Determine navigation method
Rachel Winters (8/32)
8
Configuration & Structural Design
Shelby Sullivan, Eric Weber
Major Tasks:
• Find camera and lens
• Camera field of view analysis
• Design structure (material, shape)
• Configure component positioning
• Mass budget
• Solidwork components (Solid modeling)
Rachel Winters (9/32)
9
Attitude Determination & Control
Shunsuke Hirayama, Tsutomu Hasegawa
Major Tasks:
• Determine method of attitude control
• Single vs Multiple cameras
• Determine pointing accuracy necessary
• Determine torque disturbances
Rachel Winters (10/32)
10
Power, Thermal & Communications
Aziatun Burhan, Masao Shimada,
Tomo Sugano
Major Tasks:
• Determine power needed by satellite
• Battery only vs Solar Cell + Battery
• Define thermal environment (outside and inside
sources)
• Determine thermal control method
• Determine transmission method
• Determine differential drag
• Integration for CPU
Rachel Winters (11/32)
11
Expendable Vs Recoverable
• Expendable
– Safer, no need to get close to orbiter for recapture
– Cost efficient if only 1 mission was planned
• Recoverable
– Less expensive to reuse
– We have a viable method of recovery
– It requires reasonable amounts of extra fuel
needed
– We are able to store back-ups of the images
onboard the satellite
• We chose a Recoverable design
Rachel Winters (12/32)
12
Single Vs Multiple Satellites
• Single
– Amount of extra fuel needed for plane transfers is
relatively small
– We have the ability to “see” entire shuttle with only
1 satellite
– Safer, there are less satellites flying around orbiter
• Multiple
– Less fuel is necessary
– Easier orbit to control, no plane transfer needed
• We chose a Single Satellite design
Rachel Winters (13/32)
13
Solar Cells + Battery Vs Battery
Only
• Solar cells + Battery
– Smaller battery needed
– Longer satellite lifetime
– Less massive
• Battery Only
– No need for attitude control, no solar cells to point at
sun
– Electrical storage independent of orbit
• We chose a design with Solar Cells and Battery
Rachel Winters (14/32)
14
Single Vs Multiple Cameras
• Single
– We have the ability to control attitude
– Reduces power, mass, and volume necessary for
multiple cameras
• Multiple
– Less accurate attitude sensors would be
necessary
– Larger field of view, it would be easier to point the
camera at the orbiter
• We chose a Single Camera design
Rachel Winters (15/32)
15
Design Walkthrough
• Assumptions and Requirements
– Mass restricted to 50 kg
– Volume restricted to 60x60x50 cm3
– Necessary to operate for 24 hours, power source must last
this long
– Must be able to detect damage as small as 1/4 inch (0.635
cm)
• Derived from Orbiter damage criteria
– Assumed a LEO orbit that was the same as the ISS orbit
• Interest: Orbiters and ISS use this orbit for extended periods of
time, small modifications to our design can allow us to orbit the
ISS or orbiters other than the shuttle
– Assumed our shuttle-relative orbit was within safety standards
(rmin = 118 m, rmax = 237 m)
– Relative velocity of less than 4 cm/s for retrieval
Rachel Winters (16/32)
16
Orbit Design
Created an orbit
– Accuracy of known
location/velocity
was important
– Maintain a “safe”
distance away from
the shuttle
– Remain within
camera range
Rachel Winters (17/32)
17
Orbit Design continued
• Determined orbital disturbances
– J2 disturbances
– Drag differences in Low Earth Orbit
– This determined the amount of thrust needed to
maintain desired orbit
Rachel Winters (18/32)
18
Orbit Design continued
• Plane change decided
– Used a plane change to keep the number of
satellites to 1 (Trade Study)
– Affects the amount of cold gas needed
Rachel Winters (19/32)
19
Satellite Design
• We determined the type of attitude control we
wanted: zero-momentum
– It allows us to control all three axes of rotation
– We needed to be able to point at the shuttle at all
times
– This determined the mode of control: Reaction
Wheels
• We already needed control to counteract
torque disturbances
– Aerodynamic torque
– Gravity-Gradient torque
– Solar radiation pressure torque
Rachel Winters (20/32)
20
Satellite Design continued
• To reduce the magnitude of disturbance torque, we
chose to design the center of mass to be in the
center of the satellite
• We chose to make the satellite a 50x50x50 cm cube
to facilitate rapid design
Rachel Winters (21/32)
21
Satellite Design continued
We modeled the components and satellite in Solidworks to map
out what we wanted it to look like
Rachel Winters (22/32)
22
Satellite Design continued
• Thermal control designed
– Found the temperature range of the environment
– Found the temperature tolerance of hardware
– Used the component layout to determine
necessary thermal control within satellite
Rachel Winters (23/32)
23
Satellite Design continued
• Power subsystem
– Approximated power drain with major components
– Made early approximation on battery/solar cell
requirements
– Determined number of solar cells we can support
• Amount of power solar cells could absorb
– Found power demand including all components
• Max: 87.6 W, Ave: 40.2 W
– Determined back-up battery requirements
Rachel Winters (24/32)
24
Power Distribution
Camera **
20%
ADCS
5%
Orbit Control (FCS)
5%
OBC
3%
Power management
7%
Communication
4%
Camera (idle)
3%
Thermal control **
13%
Propulsion **
40%
** non continuous operation
Rachel Winters (25/32)
25
Mass Budget
Piece
Dimensions (cm)
Mass (g)
Volume(cm^3)
Thruster x4
1.5D x 3.0
23
2.8
N2 Tank
12.0D x 28.9
4000
2251.5
Camera
15.0 x 15.0 x 4.0
960
900.0
Lens
16.0D x 35.0
1270
4889.2
Battery x2
15.0 x 7.0 x 7.0
1400
735.0
Electronics Controller
30.0 x 30.0 x 7.0
3500
6300
CPU
9.6 x 9.1 x 2.0
96
174.7
Star Tracker
5.4 x 5.4 x 15.6
600
1011.5
Reaction Wheel
11.8 x 11.8 x 11.5
1986
1272.6
Gyro
9.9 x 11.7 x 3.1
439
359.0
GPS Comp x2
14.0 x 8.9 x 11.7
450
1458.0
GPS antennas x2
3.8 x 2.9 x 1.1
20
11.6
Solar Cell Panels x5
48.0 x 48.0 x 0.5
3365
1152.0
Aluminum Panel
48.0 x 48.0 x 0.6
3110
1152.0
Satellite Structure
50.0 x 50.0 x 50.0
7434
125000.0
Mass of Hardware
Extra Parts (conservative estimate)
Rachel Winters (26/32)
Total
44.0
kg
5
kg
49.0 kg
26
Hardware determination
• Camera & Lens – MegaPlus
– Small mass; operable in
space conditions
– This determined an orbit
range to stay within
– Able to capture 39% of
shuttle at 180 m separation
• Transmitter – Zonet Wireless
LAN Adaptor
http://www.amazon.com/802-11G-Wireless-Adapt-FROM100Meters/dp/B000MN8MV4
Rachel Winters (27/32)
– Able to transmit large amounts of
data over orbit range
– Operating range up to 300 m
– Connects with USB port
– Orbiter must also be connected
to wireless network
27
Hardware continued
• Star tracker – Sun Space
– Very accurate
• within .001 degree
• 0.1 degree accuracy
needed
– Must not be exposed to
sunlight
• Gyro – Sun Space
– Adds accuracy to attitude
determination
– Part of the Reaction
Wheel system, separate
casing
Rachel Winters (28/32)
28
Hardware continued
• Reaction Wheel Assembly
– Sun Space
– Keep shuttle in field of view
– Induce up to 50 mNm of
Torque
– Includes 3 wheels with 3-axis
control
• CPU – Arcom Viper
– Provides computer processing for
• attitude and orbit control
• picture and transmission
commands
– Includes two USB ports
• Used for transmitter
• Possible picture storage
Rachel Winters (29/32)
29
Hardware continued
• CDGPS - NovAtel
– Carrier Phase Differential
GPS determines shuttlerelative location and
velocity
– Light weight
– Includes two antennas, two
receivers
Receiver
Rachel Winters (30/32)
Antenna
30
Hardware continued
• Thrusters - MOOG
– Needed to make orbital
changes
– 4 Thrusters aligned with
the center of mass
– Cold gas thruster system
• Gaseous Nitrogen
• Tank - Airhog
– Pressurized to about 21
MPa
– 1700 cm3 volume
Rachel Winters (31/32)
31
Hardware continued
• Battery - Saft
– Satellite needs power to
operate for 24 hours
– Use solar cells  minimize
battery demand
– 2 Batteries for redundancy
– 1 hour charging time,
recharges to 75%
– Lithium-ion cell batteries
• Heater
– Some devices are
temperature sensitive
– Maintains temperature of
satellite within allowable
range
Rachel Winters (32/32)
32
Configuration and Structural Design
Eric Weber
• Tasks
–
–
–
–
–
–
Preliminary hardware research
Preliminary transmission research
Materials Research
Solid Modeling
Satellite Configuration
Kept Mass Budget
Total Hours: 75
Eric Weber (1/14)
33
Configuration and Structural Design
• Design Requirements
– University Nanosat
• Must be less than 50x60x60 cm
• Must be less massive than 50 kg
– Mission specific
• Must withstand mission of 24 hours
Eric Weber (2/14)
34
Configuration and Structural Design
• Derived Requirements and Goals
– Center of mass to be within 1.5 cm of geometric
center (ADC)
– Camera and Star Tracker looking at 90º angle
from each other (ADC)
– 4 Thrusters (ADC/Orbit)
• Thrust line on the same plane
• Through center of mass
• At 90 º angles from the camera
Eric Weber (3/14)
35
Configuration and Structural Design
• Derived Requirements and Goals (cont.)
– Solar cell panels on all faces except camera face
(Power)
– Attempt to keep components away from satellite
walls (Thermal)
Eric Weber (4/14)
36
Configuration and Structural Design
• Must be less than 50x60x60 cm
50 cm
50 cm
50 cm
Eric Weber (5/14)
37
Configuration and Structural Design
Piece
Dimensions (cm)
Mass (g)
Volume(cm^3)
Thruster x4
1.5D x 3.0
23
2.8
N2 Tank
12.0D x 28.9
4000
2251.5
Camera
15.0 x 15.0 x 4.0
960
900.0
Lens
16.0D x 35.0
1270
4889.2
Battery x2
15.0 x 7.0 x 7.0
1400
735.0
Electronics Controller
30.0 x 30.0 x 7.0
3500
6300
CPU
9.6 x 9.1 x 2.0
96
174.7
Star Tracker
5.4 x 5.4 x 15.6
600
1011.5
Reaction Wheel
11.8 x 11.8 x 11.5
1986
1272.6
Gyro
9.9 x 11.7 x 3.1
439
359.0
GPS Comp x2
14.0 x 8.9 x 11.7
450
1458.0
GPS antennas x2
3.8 x 2.9 x 1.1
20
11.6
Solar Cell Panels x5
48.0 x 48.0 x 0.5
3365
1152.0
Aluminum Panel
48.0 x 48.0 x 0.6
3110
1152.0
Satellite Structure
50.0 x 50.0 x 50.0
7434
125000.0
Mass of Hardware
Extra Parts (conservative estimate)
Eric Weber (6/14)
Total
44.0
kg
5
kg
49.0 kg
38
Configuration and Structural Design
• Must withstand mission of 24 hours
– Short satellite mission time
– Satellite is to be carried up with shuttle
• Greatest force < 50 N
– Material Selection
• Aluminum 6061 T6
– Availability
– Cheap
– Adequate
Eric Weber (7/14)
39
Configuration and Structural Design
• Center of mass calculation
N
xcm 
m x
i i
i 1
M
• SolidWorks can calculate these numbers
– Iterative process to find an adequate center of mass
• Center of mass location offset from geometric center
– X = 0.29 cm
– Y = 0.15 cm
– Z = 0.01 cm
Eric Weber (8/14)
40
Configuration and Structural Design
Eric Weber (9/14)
41
Configuration and Structural Design
• Camera and Star Tracker looking at
90º angles from each other
Thrusters
• 4 Thrusters facing each other with
thrust line through the center of
mass at 90 º angles from the
camera
Eric Weber (10/14)
42
Configuration and Structural Design
• Solar cell panels on all sides except camera
side
Eric Weber (11/14)
43
Configuration and Structural Design
• Attempt to keep components away from
satellite walls (Thermal)
Eric Weber (12/14)
44
Configuration and Structural Design
• All requirements and goals were met
 Must be less than 50x60x60 cm
 Must be less massive than 50 kg
 Must withstand mission of 24 hours
 Center of mass is to be within 1.5 cm of geometric
center
 Camera and Star Tracker looking at 90º angle from
each other
 4 Thrusters facing each other with thrust line through
the center of mass at 90 º angles from the camera
 Solar cell panels on all sides except camera side
 Attempt to keep components away from satellite walls
Eric Weber (13/14)
45
Configuration and Structural Design
• Recommendations for Future Work
– Thin out aluminum frame and supports
• For lower mass
– Model small parts
• For more accurate model
Eric Weber (14/14)
46
Shelby Sullivan
Major Tasks:
Camera and Lens Determination
Picture Quality Analysis
Solidmodeling
Satellite Configuration
Hours Worked:
78
Shelby Sullivan (1/17)
47
Camera
MegaPlus II EP1600
16 Megapixel
4872 x 3248
Three sensor grades for
“demanding
applications”
Selectable 8, 10, or 12 bits/pixel
“Temperature Resistant”
construction
Shelby Sullivan (2/17)
48
Lens
Canon EF 400mm f/4 DO IS USM
•
•
•
•
Angle of view – 6.2° x 4.1°
Length – 24 cm
Mass – 2 kg
Fixed focal length
– Little to no moving parts
– Higher vibration resistance
– Higher temperature resistance
Shelby Sullivan (3/17)
49
Field of View
Shelby Sullivan (4/17)
50
Picture quality versus Distance relationship
Distance vs Pixels per meter
400mm
500
450
Pixels per meter
400
350
300
250
200
150
100
50
0
0
200
400
600
Distance (m)
800
1000
1200
51
Discovery photograph from ISS
during 2005 rendezvous
• 225 Pixels per Meter
• 3 cm gap filler
protrusion
• ~3x1 cm chip in tile
• Both points repaired
on following space
walk
• Equivalent to FF
satellite ~240m from
shuttle with current
camera/lens
configuration
52
Ceramic tile embedded in an oak
tree from the shuttle Columbia
Shelby Sullivan (8/17)
53
Same tile reduced to
250 pixels/meter
Equivalent to FF
satellite 180 meters
from shuttle
~1cm chips still
identifiable
Shelby Sullivan (9/17)
54
Full shuttle in field of view vs.
Altering the pointing angle
• Shuttle ~ 37m in length
• 250 pixels/meter requirement
• To obtain similar quality images with full shuttle in
FoV requires
9250 horizontal pixels
Equivalent to 57 mega pixel camera
*Independent of lens and distance from shuttle
Shelby Sullivan (10/17)
55
Full shuttle in field of view vs.
Altering the pointing angle
• Take multiple pictures from the same location
while altering pointing direction
• All pictures taken from same distance to allow
for fixed focus lens
• Current configuration 250 pixels/meter
obtained at 180 meters from shuttle.
Shelby Sullivan (11/17)
56
~180m From Shuttle
• ~Cross-sectional are of shuttle
– 640 m^2
• Field of View area
– 252 m^2
• ~39% of shuttle captured per photo
• Allowing for overlap
– Take 9 pictures at each location 180m from shuttle
Shelby Sullivan (12/17)
57
Imaging Sequence
First 6 of 9 pictures shown
2.8 degrees horizontal shift or
0.7 degree vertical shift
between pictures
Shelby Sullivan (13/17)
58
Pointing Accuracy Required
Horizontal Angle < Vertical Angle
Shelby Sullivan (14/17)
59
Orbital Position for Pictures
Amount of Shuttle Seen over 1 Period
70
60
% Shuttle Seen
50
40
30
%shuttle seen
180 Meters from Shuttle
20
10
0
0
50
100
150
200
250
300
350
400
1/360 of a Period
Shelby Sullivan (15/17)
60
Final Picture Information
• Distance of 180m obtained at ±37° from apex
• 250 Pixels/meter
– Allow for ~1cm imperfections
• 9 Pictures per point
– Allowing for 1.6° error in accuracy
• 4 Points per orbit
– Giving total shuttle coverage from all sides
• 16 Orbits per mission
• Total of 576 pictures
~9 Gigabytes of data
Shelby Sullivan (16/17)
61
Conclusions and Recommendations
Conclusions
• Increased image quality over previously used
methods
• Full shuttle coverage from multiple vantage
points
Recommendations
• Implement automatic focusing lens for taking
pictures at multiple distances
• Simulation of entire image taking sequence
Shelby Sullivan (17/17)
62
Kyle Tholen
• Major tasks:
– Method for Navigation
– Orbit Determination
– Estimate Delta V Requirements
for Orbit Transfers
• Hours Worked: 80
Kyle Tholen
63
Navigation
• Three Options:
– Global Positioning System (GPS)
– Differential Global Positioning System (DGPS)
– Carrier Phase Differential Global Positioning
System (CDGPS)
Kyle Tholen
64
GPS
• GPS can be used to determine the satellite’s
position and velocity
• GPS satellites transmit two signals:
– Precise Position Service (PPS)
• Very accurate
• Restricted to military applications
– Standard Position Service (SPS)
• Available for civilian use
• Not accurate enough for our application
Kyle Tholen
65
DGPS
• Much more accurate than GPS
• Requires a known, fixed position (Space
Shuttle)
• Compares the signals that satellite and
shuttle receive to eliminate errors
• In theory it can be very accurate
• In practice, sometimes only accurate to within
50 meters
Kyle Tholen
66
CDGPS
• Uses both PPS and SPS signals
• Compares the signals that the satellite and
the shuttle receive
• Most commonly used form of GPS in LEO
• Demonstrated accurate to within 5 cm
Kyle Tholen
67
GPS Conclusions
• Use CDGPS
• Use NovAtel’s Superstar II GPS card
– Supports CDGPS
– Low power consumption
Kyle Tholen
68
Orbit Determination
• Used Clohessy Wiltshire equations to analyze
orbits
• Need a minimum of two orbits to see shuttle
from all angles
• Orbits achieved through small changes in:
– Eccentricity, de = .000026
– Inclination, di = .0004
– Right Ascension, dRA = .0014
Kyle Tholen
69
Delta V Estimation
• Separation distances achieved:
– Max = 237 m
– Min = 118 m
• Six Impulses
–
–
–
–
–
–
First Impulse = .246 m/s
Second Impulse ~ 0
Third Impulse = .657m/s
Fourth Impulse ~ 0
Fifth Impulse = .295 m/s
Sixth Impulse ~ 0
Kyle Tholen
70
Rendezvous
• Need satellite to be at a distance of 37.59
meters
• Shuttle then uses robotic arm to grab satellite
• Relative speed between the two = .0333 m/s
Kyle Tholen
71
Single vs. Multiple Satellites
Parameter
Single
Multiple
Cost
Less
More
Delta V
requirements
More
Less
Safety
More
Less
Ability to perform
the mission
Capable
Capable
72
Cost
• Identical systems and satellite configuration
–
–
–
–
–
Structure
Communications
Propulsion
Power and thermal management
Mission cost nearly doubles with additional
satellite
73
Delta V Requirements
• Single Satellite
– Delta V due to orbit transfers
– Delta V for orbit maintenance
• Multiple Satellite
– Same Delta V for orbit maintenance
– No orbit transfers required
– Delta-V for transfer (.606 m/s) is small compared
to delta-V for orbit maintenance (1.36 m/s)
74
Safety and Mission Performance
•
Safety
– Fewer satellites, less chance for guidance
malfunction or shuttle contact
•
Performance
– Both configurations can perform mission
requirements
– One satellite may be able to accomplish mission
faster
75
Conclusions
• Single satellite is:
–
–
–
–
More cost effective
Slightly larger Delta V requirement
Safer
Capable of performing mission
• Single satellite is the optimal choice
76
Future Work
• Create more detailed model of orbit,
including:
– Differential drag correction impulses
– Random errors due to inaccuracy in position
determination
Kyle Tholen
77
Matt Mueller
Major Tasks:
• Effect of Earth’s oblatness on relative distance to
shuttle
• Propulsion method, selection of propulsion system
and components
• Demonstration of relative orbit
Total hours worked: 85 hrs
78
Effect Of Earth’s Oblatness
• Causes secular drift in right ascension, argument of
perigee and mean anomaly
 J2R 
3
dot   
cos i
2 2 7/2 
2  (1  e ) a 
 J 2 R  5 2
3

dot   
sin
i

2

2 2 7 / 2 
2  (1  e ) a  2

 /a

2
J2R 
3 a
2
2

Mdot  
3
sin
i

2
1

e
4  a2 1  e2 2 








79
Matt Mueller
Earth’s Oblatness Continued
• Magnitude of changes in orbital elements highly
dependent in inclination angle
6
Degrees/12 hrs
4
2
0
0
20
40
60
80
100
RA
AP
MA
-2
-4
-6
Inclination (deg)
Matt Mueller
80
Relative Separation
• Adjustment of inclination angle at ½ mission lifetime
causes changes to go to zero
• Only concerned with relative separation for 12 hr
period
  .000019
  .000051
M  .000028
• Maximum relative distance change of 3.4 m
• Too small to justify for orbit correction
81
Matt Mueller
Propulsion System Selection
Requirements
• Thrust rating capable of
providing short burn times
• Small and light weight
MOOG 58-118 Cold Gas Thruster
• Specific impulse = 60 sec
• Thrust = 3.6 N
• Weight .023 kg
82
Matt Mueller
Thruster Selection
• Burn times dependent on thrust, mass of
satellite and delta-V requirement
Satellite specifications
F
V 
T
M
M = 50 kg
delta-V max = .2119
m/s
• Delta-T = 2.94 sec
• MOOG model 58-118 fulfills all requirements
83
Matt Mueller
Gas Tank Sizing
Requirements
• Fairly small size to fit nanosat requirements
• High pressure capability
• Thermal environment from -82.15 to 77.2 degrees
Celsius
Sizing
• Pressure increases with temperature, T = 77.2 deg
• Mass of gas, tank operating pressure and
temperature determine size
• N2 gas gives reasonable specific impulse and is inert
84
Matt Mueller
Tanks sizing continued
For orbit transfer and drag
Specifications
• Delta-m = .167 kg
m = 50 kg
•Assume S.F. = 2
•Total mass = .333 kg
• At 20.68 MPa, V = 1700 cm^3
• Volume requirement met to
fulfill requirements
Specific impulse = 60 sec
Total delta-V = 1.97 m/s
m
 1 e
m
 V
I sp g 0
m
PV 
RT
M
85
Matt Mueller
Orbit Modeling
Formation Flying Simulation
• Supplied code allowed for
3D relative orbit modeling
• Varied orbital elements,
relative ellipse parameters,
mass, area, thrust, etc.
• Visualization of relative
motion, orbit track over time,
estimation of burn times for
orbit transfer and orbit
keeping.
86
Matt Mueller
Attitude Accuracy and Orbit Control
• Accuracy error, α
creates undesired
delta-v
• α = .0061 deg
• Delta-v = .003134
mm/s
• Analysis using Hill’s
equations provide
impact on relative
motion
87
Matt Mueller
Attitude Accuracy and Orbit Control
-4
x 10
3
2
x (radial) [km]
1
0
-1
-2
-3
-2.0004
-2.0003
-2.0002
-2.0001
y (along-track) [km]
-2
-1.9999
• Causes 25 cm drift in along-track direction over 6 orbits
• Negligible, total mission lifetime drift of < 65 cm
• 10x less accurate sensor causes 38 m drift
88
Matt Mueller
Conclusions and Recommendations
Conclusions
• No need to make orbit keeping maneuvers due to J2
• Cold gas thruster selected fulfills requirements,
allows for short burn times
• Relative elliptical orbit verified
• Estimated drift due to inaccuracies in attitude
determination negligible
Future Work
• Select tank material suitable for thermal environment
• Continue simulation and determination of burn times
necessary to compensate for differential drag.
89
Matt Mueller
Major tasks
Tsutomu Hasegawa
(Shunsuke Hirayama)
•
•
•
•
•
Determine method of attitude control
Single VS Multiple cameras
Determine pointing accuracy necessary
Determine torque disturbances
Determine equipments to keep the shuttle in
our Field of View.
• Total hours worked: 85 hours each
90
Tsutomu Hasegawa
Why Zero-momentum?
Type
D of C
Gravity inclination
0
cost
Accuracy
Low
Low
Spin
Dual Spin
Attitude
Change
Heat generation
Can't change
Small
Other
None
Difficult
1
Nutation
Bias Momentum
Weak from disturbance
For small
Dumper
Large
Controled bias
Components
2
momentum
3
High
High
Easy
Mass Wheel
communication Sat.
Bias+Roll
Reaction
Weak from disturbance. For
Wheel
communication Sat. Don't use
Momentum Bias+Roll/Yaw LEO
Reaction Strong from disturbance
For Large Sat. LEO
Zero momentum
Wheel
is OK
Zero momentum: Each axis is independent. It is easy to control.
91
Tsutomu Hasegawa
Attitude Determination
• Fixed
• Pointing the same direction
x
y
Tsutomu Hasegawa
z
92
Pointing Direction
Put conditions from orbit control team into Joe’s code. 93
Tsutomu Hasegawa
Pointing Direction
The Satellite orbit
Pointing directions
Pointing error
The Center of the shuttle
• There are pointing errors
Tsutomu Hasegawa
94
Pointing Error
Worst case: Maximum pointing error is about 33 degrees. 95
Tsutomu Hasegawa
Multiple Camera
Case when  is larger than 
Capture angle, 
Field of View, 
Length of the shuttle
(The longest)
r
• To get large  , we should use multiple cameras.
• It is easy to control, and need more space for camera.
Especially, when     33
Tsutomu Hasegawa
96
Single Camera
Case when  is smaller than 
Field of View, 
Capture angle, 
r
Worst case: r  220[m]
  18.6   6.195
• Because of small  , we have to change the direction.
• Need more accurate sensors.
97
Tsutomu Hasegawa
Multi VS Single camera
• Multiple camera
Easy to control
Need space to put camera
• Single camera
Need to control to point the shuttle.
98
Tsutomu Hasegawa
Major tasks
Shunsuke Hirayama
(Tsutomu Hasegawa)
•
•
•
•
•
Determine method of attitude control
Single VS Multiple cameras
Determine pointing accuracy necessary
Determine torque disturbances
Determine equipments to keep the shuttle in
our Field of View.
• Total hours worked: 85 hours each
Shunsuke Hirayama
99
Disturbance
Aerodynamic torque:
Taero  rcp  Faero
Solar Radiation Pressure Torque: Tsolar  rcp  1  K A
Gravity-Gradient Torque:
Magnetic torque:
Is
c
Tgravity  3n 2 rˆ  I  rˆ
Tmagnetic  M  B
Ttotal,dist  Taero  Tgravity  Tsolar  Tmagnetic
 0.0373 


3
 10   0.5070  Nm
 0.4521 


Shunsuke Hirayama
100
Changing Orbit Requirement
from 1 to 2
90
T 
Orbit 1: Blue
Orbit 2: Red
90 min
4
Pointing thrust
req
90

t
deg/s
90

90 min
4
 0.0666 deg/s
101
Shunsuke Hirayama
Pointing requirement
Pointing camera:  req
Pointing
requirement :
 req  15 
shuttle
 req
FOV :   6.20
6.20

5
 1.24
102
Shunsuke Hirayama
-Simulation for attitude controlQuaternion
Rate limit : 0.015 rad/s
Control angle : 1.24 deg
103
Shunsuke Hirayama
Angular error input
Stable after 100 sec
104
Shunsuke Hirayama
Reaction wheel
Tcon. max  4 m Nm
TRW . max  50 m Nm
ReactionWheelsR03
50mNm
R03
Shunsuke Hirayama
(Sun space)
105
Attitude sensor
Accuracy of sensor :  req  1

10 req
 0.124
Star tracker: star tracker R01 ( Sun space)
 sensor  0.001
 sensor   req
106
Shunsuke Hirayama
Conclusion
• Fixing pointing error : Satisfied
• Pointing accuracy necessary : Satisfied
• Equipments to keep the shuttle in our FOV are determined
• We need to survey what to do when the side with star
tracker accidentally see the sun directly.
• We need to simulate combined cases.
• We need to concern electric power.
107
Shunsuke Hirayama
Systems Integration & Management
Matthew S. Whitten
Major Tasks:
– Recoverable v. expendable trade study
– Research and design recovery method
Total Hours: 79 hours
108
Matthew Whitten (1/12)
Trade Study
Parameter
Recoverable
Expendable
Single Mission Cost
More
Less
Multiple Mission Cost
Less
More
Quality/Accuracy
More
Less
Design Difficulty
More
Less
Safety
Less
More
109
Matthew Whitten (2/12)
Trade Study
• Single Mission v. Multiple Missions
– A single mission satellite will cost less as it does
not need the final transition back to the shuttle.
– A recoverable satellite is ideal for multiple
missions because the cost of two satellites would
be more than the extra cost (such as operating the
robotic arm to capture the satellite at the end of its
mission) needed to make the satellite recoverable.
110
Matthew Whitten (3/12)
Trade Study
• Components and Quality/Accuracy
– Higher quality and more accurate components can
be placed in a recoverable satellite.
– Expendable depends on satellite sending images
back to shuttle.
111
Matthew Whitten (4/12)
Trade Study
• Safety of Design
– The design difficulty of a recoverable satellite is
greater than that of an expendable due to the
capture phase.
– Other than the possibility of additional space
debris, the expendable satellite is a safer design
than recoverable because it does not approach
the shuttle’s vicinity.
112
Matthew Whitten (5/12)
Trade Study Conclusion
• Formation flying satellite assumptions
– Used defensively
– Sustain little damage
– Provide a service for any re-entering vehicle
• Recoverable satellite reasoning
– Mission requires high quality components
– Multiple missions are likely
113
Matthew Whitten (6/12)
Recovery Method
Requirements
– Shuttle is 37 m in length by
17.3 m in height
– Capture orbit must be
outside the shuttle’s radius
with certainty to avoid
collision
– Satellite must be moving
with a velocity of zero
relative to the shuttle
– Must have correct attitude
control when entering
capture sequence
114
Matthew Whitten (7/12)
Recovery Method Continued
• Calculated capture orbit radius
Rc = L/2 + H/2 + 10 m (error) = 37.59 m
• Parameters for capture orbit (with assistance of
ADCS and MATLAB code):
e = 0.0021785
Right Ascension (deg) = 124.747
Argument of perigee (deg) = 127.754
True Anomaly (deg) = 180
• Worst case relative recovery velocity
V (with 2% error) = 0.036 m/s = 3.6 cm/s
[Calculated with distance and rotation rate geometry and
compared with V from MATLAB code of 3.3 cm/s]
115
Matthew Whitten (8/12)
Recapture Method
Continued
Length
45 feet
Diameter
15 in
Weight
911 lb
Number
of joints
Six joints (two shoulder joints, one
elbow joint, and three wrist
joints)
Max
handling
capacity
266 tons (in space)
Max
velocity of
end of arm
When the arm is not gripping
anything : 60 cm/sec
When the arm is gripping an
object : 6 cm/sec
Max
rotational
speed
Approx. 5 degree/sec
Boom
material
Graphite-epoxy compound
Matthew Whitten (9/12)
Robotic arm
116
Recovery Method Continued
End Effect
117
Matthew Whitten (10/12)
Recover Method Continued
118
Matthew Whitten (11/12)
Conclusions
• It is most cost effective to have recoverable
satellites
• Using an end effect on the robotic arm with an
imagery boom will be effective in capturing the
satellite
Future Work
• Simulate capture
• Automate end effect to not depend on human
control
119
Matthew Whitten (12/12)
COMM Subgroup
Tomo Sugano
Major Tasks:
– Started out by helping Relative Orbit team with
differential delta-V requirement and GN2 mass
– Communication Scheme of the mission
– System Integration (on-board flight computer)
Hours Worked: 60
Tomo Sugano (1/16)
120
Differential Drag Compensation
• Ballistic coefficients different between satellite
and shuttle
• Difference in drag forces (i.e. acceleration)
exists.
• (B-Coeff. of Shuttle) >> (B-Coeff. of Satellite)
• Satellite must expend Delta-V to keep up with
space shuttle.
Tomo Sugano (2/16)
121
Computations
• Atmospheric drag acceleration (Da):
• Drag acceleration difference between the two spacecraft:
STS: S = 64.1 m2, CD = 2.0, m = 104,000 kg (orbiter typical)
Sat: S = 0.385 m2 (nominal), CD = 3.0 (worst case), m = 50 kg
Tomo Sugano (3/16)
122
Computations (cont’d)
• Orbiter speed (assuming circular orbit)
• Definition of Delta-V (or specific impulse)
• Mass expenditure of propellant (i.e. GN2 cold gas)
Tomo Sugano (4/16)
123
Results
• Using Isp = 65 sec; assume 50 kg for satellite weight
• Conclusions
- At the typical 300 km LEO, Delta-V for 1 day mission is
1.36 m/s
- Satellite will need at least 107 grams of GN2 to
compensate drag
- In addition to this Delta-V, we have orbit transfer Delta-V
and ADCS Delta-V.
 How we determined tank size.
Tomo Sugano (5/16)
124
COMM: Overview
Satellite must be able to:
Satellite
Wireless LAN
Transceiver
• Receive data from
the space shuttle
Tx
• Transmit data to the
space shuttle
Space Shuttle
Orbiter
Tomo Sugano (6/16)
Satellite
Computer
Rx
Receive
Commands
Transmit
Photos
125
COMM: Operational Schematic
Satellite
Image
Buffer
Space Shuttle
Tx
Downlink
Uplink
Rx
FCS
CPU
CPU
Rx
Archival
Memory
LAN
Uplink
Downlink
Tx
LAN
RAM
Data Flow
Tomo Sugano (7/16)
126
COMM: Communication Devices
Local Area Network (LAN)
Protocol is utilized.
Satellite
- A Zonet light-weight wireless
LAN transceiver
Space Shuttle
- An integrated wireless LAN
transceiver of a Panasonic
Toughbook
Tomo Sugano (8/16)
127
COMM: Zonet Wireless LAN USB Adapter
•
•
•
•
•
•
•
Model: Zonet ZEW2502 802.11g USB Adapter
Dimensions: 93.5 x 27.5 x 10 mm
Mass: 32.0 grams
Transfer Rate: 54Mbps (nominal)
Uplink (Receive): 260 mW
Downlink (Transmit): 350 mW
Frequency:
- 2.412GHz ~ 2.4835GHz
- S-band (space approved, open band)
Tomo Sugano (9/16)
128
COMM: Antenna Modification
• Zonet LAN adapter uses “Microstrip antenna.”
• Compact and reliable for short-distance
communication.
• Directly attach antenna on the face of satellite
Tomo Sugano (10/16)
129
COMM: Image Transfer
• Estimated size of each image: 14-16 Mb
• Our LAN protocol and devices allow nominal
54Mbps with pier-to-pier method.
• Virtually no cause of appreciable noise in
space between satellite and space shuttle
• Transfer time of each image: less than a few
seconds.
• Image buffer need not to be large
- A few GB USB jump drive recommended.
Tomo Sugano (11/16)
130
Selection of System Integration CPU
• System integration is core of the FCS
VIPER CPU board top view
 Selection of flight computer CPU
• Arcom VIPER 400 MHz CPU found viable
• VIPER is suitable because of its
- Light weight, 96 grams
- Operable temperature range
-40 C to + 85 C
- Windows Embedded feature
 easy to program
- Computation speed, 400 MHz
- Memory capacity, up to 64MB of SDRAM
- 5 serial ports
- USB ports enables utilization of Zonet LAN adapter & jump drive
Tomo Sugano (12/16)
131
Control System Integration - Overview
•
•
•
•
ADCS and Orbit Rectification Computations
Execute ADCS and Orbit Rectification
In-flight image capture handling
Data archiving
Tomo Sugano (13/16)
132
System Integration – Overall Scheme
Flight Computer
POWER
FCS
COMM
ADCS Output
Sensors
Orbit Control Output
Other Outputs
Tomo Sugano (14/16)
133
System Integration – Data Connections
Star Tracker
Spin Wheels
Serial 2
Thrusters
Serial 3
Gyro
Serial 4
GPS
Serial 5
Flight Computer
Serial 1
FCS
PWR Flow
Digital I/O
POWER
Tomo Sugano (15/16)
Camera
USB 1
Wireless
LAN
USB 2
Archival
Memory
134
Conclusions & Recommendations
• Computed GN2 mass for flight
• Transmission noise is not an issue
• Wireless LAN is an efficient method
to meet our needs
• CPU has integrated our hardware
• Recommend that we monitor
remaining GN2 and implement inflight HMCS
Tomo Sugano (16/16)
135
Electrical Power System
Tasks:
Aziatun Burhan
- Solar cells & battery vs non rechargeable battery trade study
- Update power budget
- Solar array & battery sizing
- Selection of solar cell & battery
- Selection of power management components
Hours: 70 hours
136
Overview
Mission requirement:
• EPS must be able to provide enough power for satellite to operate for
24 hours.
Design goal requirements throughout mission operation:
•
•
•
•
Generate power using solar cells
Store power using batteries
Regulate power for satellite operation
Distribute power to satellite components & hardware
137
Power Budget
Subsystem
Power consumption
(W)
Total Power
(W)
Total Power (with
20% margin)
Continuous operation
ADCS
4
Orbit Control (FCS)
4
OBC
2
Power management
5
Communication
Camera (idle)
3.05
3
21.05
25.26
55
66
73.05
87.66
Non-continuous operation
Thermal control
10
Propulsion
30
Camera
15
Total (maximum)
138
Power Consumption
Camera **
20%
ADCS
5%
Orbit Control (FCS)
5%
•Continuous power per one
orbit period: 25 W
•Maximum power per one
orbit period: 87.6 W
OBC
3%
Power management
7%
Communication
4%
Camera (idle)
3%
•Average power per one orbit
period: 40.2W
•Average power during
daylight: 36 W
•Average power during
eclipse: 46.5 W
Thermal control **
13%
Propulsion **
40%
** non continuous operation
139
Mission Power Profile
Power profile for 24 hours period
70
1 orbit cycle
burn
Power (W)
Image capture
25
shadow
Time (s)
140
Power Load Profile
50W
65W
In the sun
Image capture
p 40W
o
w
e
r
55W
Thermal
control
In the sun
Burn
35W
25W
5400s
•Typical orbit
•Period: 90 min
•No burn for orbit
transfer
5400s
•Launch orbit
•Period: 90 min
•1st & 2nd burn is taking place
141
•No picture is taken
Power Load Profile
65W
In the sun
55W
In the sun
burn
p
o
w
e
r
35W
•8th orbit cycle
•3rd & 4th burn is taking place
•16th orbit cycle
•5th & 6th burn is taking place
142
Solar cells vs battery
•
This trade study examines the option of using only non-rechargeable
battery versus solar cells & rechargeable battery system as satellite’s
power source.
•
Factors that effect trade study:
- Total power consumption and power profile during the mission
- Mission life of satellite is at least 24 hours
- Mass and area constraints that come from NANOsat requirement.
- A source power with small mass & size is desirable
- Choice of orbit and type of attitude control.
-LEO has the most frequent eclipse.
-Altitude and inclination affect the orbital period and shadow time.
- Operating environment
- Temperature
143
Solar cells vs Battery
Factors
Solar cell & battery
Non rechargeable battery
Mass
Less
More
Power/mass
High
Low
Mission length
Weeks-months
Few hours-one/two days
Attitude/Orbit control
Dependant
Independent
Effect of Failure
No/less power during
eclipse
No/less power throughout
mission
Thermal control
Less dependant
Dependant
Cost/power
Low
High
Structural support
None for body mounted
solar array
None
144
Solar cells
•Ultra triple junction Gallium Arsenide solar cell
• 28.0 % BOL efficiency
•2.31 V, 16.3 mA/cm² ( ~0.96 W/ cell)
•2.3g
Customized size:
3.69 cm
Area per cell: ~ 25cm²
6.85cm
145
Solar cells layout & assembly
Sides
• 5 identical solar panels
• Power source: Direct sunlight at 45 ° angle
• 91 solar cells per side (UTJ GaAs)
- 13 solar cells per string (column) : 30 V
- 7 strings (columns) : 2.85 A
• 40.7 Watt per side in the sun
• Dimension : 48 cm x 48 cm
• Area: 2275 cm²
• Total power: 81.4 Watt
Assumption:
•2 sides are illuminated
•Solar flux is at constant value of 1353 W/m^2
•Thermal efficiency, η= 0.74 (assuming hottest temperature =80°C)
146
Rechargeable Batteries
• Saft MPS 176065 Lithium-ion cells
• 8 cells in series in a battery box
• Capacity: 5.8 Ah
• Mean voltage: 3.6 V
• Battery mass: 1.4 kg (including casing)
• Maximum DOD: 75% for <500 cycles
• Charging method: Constant Voltage-constant
current + balancing
Energy storage requirements:
Peak power load: ~87.5 W
Shadow time: 36 min (maximum)
Charging time: 1 hour
Charge/Discharge cycle / day: 16 cycles
• Space qualified
• 2 battery boxes (for redundant operation with
one unit failed)
Required capacity: 5.7 Ah for 30% DOD
147
Power Output
Power per orbit
100
90
80
ideal power from
solar panel
Power (W)
70
60
50
power from
battery discharge
at C/2
40
30
20
10
0
0
0.2
0.4
0.6
0.8
1
1.2
1.4
time (hour)
148
Power Management & Distribution
•SmallSat power management electronics
•28V unregulated; MPPT; Modular & Scalable
from 20W to 300W
•Consist of 3 main elements:
- Battery Charge Regulator (BCR)
- Power Conditioning Module (PCM)
- Power Distribution Module (PDM)
149
Power Distribution Design
ADCS
DC-DC
Step down
converter
Power
Bus
(~28 V)
COMM
FCS
OBC
Thermal
5V supply line (200-700 mA)
DC-DC
Step down
converter 12 V supply line (170 - 1250mA)
Propulsion
Payload
(Camera)
Power management
150
Conclusion
•
•
•
EPS can provide enough power for satellite operation.
With solar cells & battery as power source, satellite mission can be
extended to few days or weeks.
Because satellite is recoverable, it is possible to use the satellite a
couple of times.
Recommendation/Future work
• Estimate power from albedo & radiation of the Earth energy.
• Optimize the design for power supply
- solar cells configuration
- battery type & sizing
• Analysis on different types of power management & distribution
modules.
151
Final Design Review
Thermal Control Subsystem
Major task:
Thermal analysis to maintain all the components of
the satellite within the respective temperature limits
by passive/active control
Hours worked:
200 hours
Masao A. Shimada
152
Mission Requirements
Mission components
* Battery is the thermally critical component.
+15
Temp. Limits Range
margin
Permissible Temp. Range
-15C margin
Temp. Limits Range
for critical component
-5
margin
Purpose of Thermal Control:
To maintain the temperature limits range of each mission component during 24-hour mission
153
Space Thermal Environment
•
•
•
•
•
•
Qs : Direct radiation from the Sun
Qe : Radiation from the Earth
Qa : Solar radiation reflected back by the earth (Albedo)
Qi : Heat generation
Governing Equation
dT
Qps : Radiation to Space
m Cp 
 Qs  Qe  Qa  Qi  Q ps  Q pe
dt
Qpe : Radiation to the Earth
Qps
Qa
Qs
Qpe
Qe
Qi
Earth pic: http://palimpsest.typepad.com/frogsandravens/pictures/earth.jpg
154
Approximated ISS Circular Orbit Model
Attitude angle
 s  59.57

Orbit plane
Satellite

 ISS  51.984
(1)
Shadow
0  0

Bottom surface
o
Earth
Circular orbit
Important constants
Solar Plane
Altitude
335 km
Period
91.22 min
Time in shadow *
30.19 min
Inclination (ISS)
51.984 deg
* Value for a period
155
Thermal Analysis
[1] Steady-State analysis for the worst cases
[2] First-order dynamic analysis
[3] Node analysis for the worst cases
I used MATLAB/Simulink® to do the simulations
156
[1] Steady-State Analysis for Worst Cases (1/2)
Assumptions
1)
Spherical satellite model of the same surface area
3)
Constants
4)
Parameters
Surface Area = 1.5 m^2
Same surface Area
Surface Area = 1.5 m^2
2)
No Radiation from the satellite to the Earth
Qpe = 0
We can change Solar absorptivity and IR emissivity
of the surface. Use the values of GaAs solar cells
157
[1] Steady-State Analysis for Worst Cases (2/2)
Governing Equation
m Cp 
dT
 Qs  Qe  Qa  Qi  Q ps  0
dt
Qps  Fps As (T 4  Tsp4 )
Solve for T
Results
 T  327 K  53.85C
  max
Tmin  211.7 K  61.48C
T  115.33
•Temperature difference is more than 100
•The result does not satisfy the requirement
For further analysis, consider Passive Contorol
carefully. If it is not enough, consider Active Control.
158
[2] First-order Dynamic Analysis (1/3)
Assumptions
1)
2)
3)
Spherical Satellite Model of the same surface area of our CubeSat
No Albedo; Qa = 0
Constants
Parameters
* Passive Control
* Active Control
Heater generating 10W during in shadow
159
[2] First-order Dynamic Analysis (2/3)
Results
m Cp 
dT
 Qs  Qe  Qi  Q ps  Q pe
dt
One Period
(T=91.22min)
 s  59.57
T [K]
Satellite

Shadow
Qs[W]
0  0
Earth
t=0
Qe[W]
Qi[W]
Tmax = 290K=16.9
Tmin = 277K=3.9
Mission required is satisfied.
160
[2] First-order Dynamic Analysis (3/3)
24 hours
Temperature [K]
24-hour (86,400-sec) profile
Qs: Yellow
Qe: Purple
Qi: Blue
[W]
161
[3] Node Analysis for the Worst Cases (1/8)
Satellite Model
Assume no width for each surface
x
y
z
Surface 2 always points downward.
162
[3] Node Analysis for the Worst Cases (2/8)
Thermal Equilibrium Equation
dTi
mi  C p 
  i Ai S i   i Ai I e Fpei  ai Ai SFsli  Qi
dt
Qs,i
Qe,i
Qa,i
4
4
  iAi (Ti 4  Tspace
) Fpsi   iAi (Ti 4  Tearth
) Fpei
Qps,i
n
n 1
j 1
j 1
Qpe,i
  Kij (Ti  T j )    i j Ai Fij   (Ti 4  T j4 )
Conduction between Node i and Node j
Radiation between Node i and Node j
163
[3] Node Analysis for the Worst Cases (3/8)
Worst-case Hot
  135.213
Worst Cases
  6.984
Worst-case COLD
s
Shadow
s
Unit: [W]
Worst-case HOT
Worst-case Cold
 0
 0
Unit: [W]
Worst-case Hot
  224.787
  6.984
164
Earth pic: http://www.bc.edu/schools/cas/geo/meta-elements/jpg/new_earth.gif
[3] Node Analysis for the Worst Cases (4/8)
Direct Solar Flux (Worst-case HOT):
QS 3
198.2W

0
QS 4
253.6W

0
QS1 , QS 2 , QS 5 , QS 6 , QS 7 , QS8

0

2
3
 s  1.04rad
Radiation from the Earth & Albedo:
I calculated the View Factors by T.C.Bannister’s approximation
165
[3] Node Analysis for the Worst Cases (5/8)
Passive Thermal Control:
•Inside of the satellite and all components are painted with L-300 (Black)
•Conductivity between surfaces : K=0.06
•Area of radiator: Qi /QR = 0.35m^2
Active Thermal Control:
Heater 10W for BATTERY
166
[3] Node Analysis for the Worst Cases (6/8)
8-Node Simulink Model
167
[3] Node Analysis for the Worst Cases (7/8)
Results
Worst-case COLD
Worst-case HOT
168
[3] Node Analysis for the Worst Cases (8/8)
Simulation Results
Solid Model
Mission Requirements
Permissible Temp. Range of each component
is in the temperature range from the simulation.
Mission Requirements is satisfied.
Theoretically, the mission is thermally in safe.
169
Conclusion
• Major task (Thermal analysis to maintain all
the components of the satellite within the
respective temperature limits by
passive/active control) has been successfully
achieved.
However, the results I got are all from
“simulation” so I need to do physical
experiments for more trusted results.
170
Demonstration
171
Electrical Discharge for Orbit
State of Charge of Battery per orbit
80
70
% State of Charge
60
50
charge at 1C
40
discharge at C/2
30
20
10
0
0
0.2
0.4
0.6
0.8
1
1.2
1.4
Time (hours)
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Recommendations for Future Work
• Reanalyze star tracker positioning
• Model small components
• Further simulation
–
–
–
–
Capture
Picture taking
Impulses
Random disturbances
• Add redundancy systems/collision avoidance
to orbit determination and control to allow for
more autonomous operation
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FMEA
Failure Mode
Electrical/wiring failure
Unable to recover
satellite
Failure of star tracker
(one attitude sensor)
Dirt or ice on camera
lens
Error in focus of lens
Collision with Orbiter
Likelihood
Solution
Medium
Recover or abandon,
Medium
Abandon, satellite orbit
depending on severity
will decay
Low – Med
Backup sensor (gyro)
Low – Med
Protective sleeve on lens,
until recovery is made
debris check pre-launch
Low
Computerized check
Very Low
Add failsafe to propel
prior to shuttle launch
satellite away from
shuttle
Coronal activity on sun
destroys satellite
Very Low
This is known in advance,
plan missions around
174
Thank you!
For all your time and assistance.
Mr. Surka
Joe Mueller
Kit Ruzicka
175
Questions? Comments?
Formation Flying
176