University of Minnesota
Senior Design II
Nanosat-V
Final Design Review
6 May 2008
Minneapolis, MN
1
Project Objective
• The aim of this project is to
perform and validate thermal,
structural and vibrational
analyses on the Nanosat-5
satellite.
• The tests will ensure that the
vehicle is capable of
withstanding loads, vibrations
and temperatures, as specified
by the University Nanosat
Program.
2
Thermal Analysis
(THRM)
Subsystem Overview
Thermal Analysis Team
David Hauth
Chuck Hisamoto
Michael Legatt
3
Objectives of Thermal Analysis
• Assemble list of material
properties, temperature
critical component profiles
• Provide thermal models of
Goldeneye with nodes for
each of the temperature
critical components onboard
• Determine hot case and cold
case thermal boundary
conditions
• Determine temperature
history for each temperature
critical component
Component Box Placement
• 3 component boxes
Battery Box
– 2 for electrical components
• GPS Receiver, Radios, etc.
– 1 dedicated for batteries
• Strict requirements for
coatings and narrower
allowable temperature range
• IMU
• Flight Computer
Component Boxes
IMU
Flight Computer
5
Thermal Analysis
(THRM)
David Hauth
6
Theory
• Conventional heat transfer through three modes
– Conduction
– Convection
– Radiation/Re-Radiation
• Most significant means of transferring energy to spacecraft
• Sources:
– Solar Radiation
» Sun radiates at black body temperature of 5777K
» Mean flux of 1367 W/m^2
– Reflected Solar Radiation (Albedo)
» Reflected and absorbed light accounts for 100% of energy received from sun
» Dependent on ground cover
» Goldeneye uses a table of average albedo for every 10 degrees of latitude
– Earth IR Radiation
»
»
»
»
Thermal equilibrium requires radiating energy equal to the amount absorbed
Higher temperature bodies emit shorter wavelengths of energy
Earth re-emits energy in the IR spectrum
Goldeneye uses a table of average IR fluxes for every 10 degrees of latitude
Analysis Input: Material Properties
–
Alodine Aluminum (6061 T6)
• Thermal conductivity:
• Specific Heat:
• Absorptivity/emissivity:
Solar:
IR:
167 W/m2
896 J/kg-K
.35
0.1
– Emcore Triple Junction GaAs Solar Cells
• Annealed at 200 deg C
• Absorptivity/emissivity:
Solar:
IR:
.92
.89
– Nusil CV10-2568 Controlled Volatility RTV Ablative Silicone
Adhesive
• Operating Temperature Range (deg C): -115 to 240
8
Internal Power Generating Components
Thermal Analysis Methodology
10
Thermal Analysis
(THRM)
Michael Legatt
11
Hot/Cold Orbits
•
•
Which orbit is hottest, coldest?
Heat Loads
– Solar Flux
– Cosmic Microwave Background
Radiation
– Internal Power Generation/Dissipation
•
–Earth Albedo
–Earth Infrared
Use Beta angle
12
Beta Angle
Solar Eclipse begins at Betastar
13
Hot Case
Occurs at:
-Beta=Beta-star
-Lowest altitude=250km
Cold Case
Occurs at:
-Beta=0
-Highest altitude=1000 km
14
Thermal Boundary Conditions
For each satellite face, MatLab/Simulink provides:
• Earth IR flux and view factor
• Earth Albedo flux and view factor
• View Factor to Space
MatLab Code Assumptions
–
–
–
–
Fluxes are date/time, attitude, altitude, orbital position
Earth Albedo, Earth IR latitude dependent
Input time, RAAN, inclination, and altitude, attitude
Solar Flux: 1327 – 1414 Watts/m2
15
Meshing Conditions
• ANSYS auto-generates mesh based on input of element sizes
– ANSYS picks element geometry type: octahedral (cube) or
tetrahedral (pyramid)
• Mesh size (approximate): ~1.0 cm
• ~760,000 Nodes
• Meshing Refinement
– ~5 million nodes
16
Thermal Analysis
(THRM)
Chuck Hisamoto
17
Temperature Critical Components
Component
Operating Temperature
[deg Celsius]
Storage
Temperature
[deg Celsius]
RTD Computer
-20 to 70
-55 to 125
NovAtel GPS Receiver
-40 to 85
-40 to 95
Kenwood TH-D7A radios
-20 to 60
N/A
SA-60C GPS antennas
-40 to 85
-50 to 90
0 to 40
-30 to 50
American Power D150-15/5 power
supply
-25 to 85
-40 to 125
HG1700 Inertial Measurement Unit
-30 to 60
-45 to 80
HMR2300 Three Axis
Magnetometer
-40 to 85
-55 to 125
Sanyo N-4000DRL batteries
Worst Hot case, Sun side
Allowable Temperature Range:
-115 to 240 deg C
Cells Annealed at
200 deg C
Hot case, bottom
Allowable Temperature Range: -115 to 240 deg C
Hot case, warmer near Standoffs
Allowable Temperature Range: -115 to 240 deg C
Hot case, Isogrids, Standoffs
Allowable Temperature Range: -115 to 240 deg C
Hot case, Battery Box
Allowable
Temperature Range:
0 to 40 deg C
Hot case, Component Box (Radio)
Allowable
Temperature Range:
-20 to 60 deg C
Hot case, Inertial Measurement Unit
Allowable
Temperature Range:
-30 to 60 deg C
Thermal Performance – Hot Case
Component
Satellite Solar
Panels/Cells
Actual Temperature Range Allowable Temperature Range Pass/
(deg C)
(deg C)
Fail
Min
Max
Min
Max
-22.41
122.76
-115
Cells
Annealed:
240
Pass
200
Battery Box
21.588
23.494
0
40
Pass
Component Box
(ADNCS, GPS, etc)
28.952
33.808
-40
85
Pass
Component Box
(Radios)
50.749
55.065
-20
60
Pass
Flight Computer
55.281
58.769
-20
70
Pass
IMU
44.647
46.555
-30
60
Pass
26
Thermal Performance – Cold Case
Allowable Temperature Range: -115 to 240 deg C
27
Cold case, hot face/cold face
Allowable Temperature Range: -115 to 240 deg C
28
Cold case, Isogrids/standoffs
Allowable Temperature Range: -115 to 240 deg C
29
Cold case, Component Box (Radios)
Allowable Temperature Range: -20 to 60 deg C
30
Cold case, Battery Box
Allowable Temperature Range: -30 to 60 deg C
31
Thermal Performance – Cold Case
Component
Satellite Solar
Panels/Cells
Actual Temperature Range Allowable Temperature Range Pass/
(deg C)
(deg C)
Fail
Min
Max
Min
Max
-35.329
29.645
-115
Cells
Annealed:
240
Pass
200
Battery Box
-19.982
-19.071
-30
50
Pass
Component Box
(ADNCS, GPS, etc)
-14.388
-12.524
-40
85
Pass
Component Box
(Radios)
-20.045
-17.584
-20
60
Fail
Flight Computer
-14.949
-14.442
-55
70
Pass
IMU
-17.539
-17.003
-40
85
Pass
32
Design Conclusions
Hot Case
Cold Case
•All temperature critical
components survive orbit
within operating ranges
•Heat accumulated on “hot
side”
-Satellite slow spin
maneuver
-Addition/changes to
coatings
•Radios component box is
slightly out of storage
temperature range.
-Need for heaters
-Small generation
needed
•All other components
survive within range
33
Acknowledgements
•Minnesota Supercomputing Institute
-H. Birali Runesha, PhD., Director of Scientific Computing and
Applications
- Ravishankar Chityala, PhD., Scientific Development and
Visualization Laboratory
- Nancy Rowe, Scientific Visualization Consultant
•Tom Rolfer, Honeywell International Inc.
•Gary Sandlass, MTS Systems Corporation
34
Supporting Slides Follow
35
36
References
• Bitzer, Tom. Honeycomb Technology. 1997.
• Curtis, Howard. Orbital Mechanics for Engineering Students. 2005.
• Gilmore, David (editor). Spacecraft Thermal Control Handbook. Vol.I.
2002.
• Griffin, Michael and French, James. Space Vehicle Design. 2nd ed. 2004.
• Kaminski, Deborah and Jensen, Michael. Introduction to Thermal and
Fluids Engineering. 2005.
• Modest, Michael. Radiative Heat Transfer. 2nd ed. 2003.
37
Supporting Slides-Task Breakdown
•
Selection of satellite structure geometry, materials, coating and isogrid patterns.
•
Design/modifications of body geometry 100% Complete
•
Design component locations/mounting 100%
•
Design torque coil mounting 100%
•
Body and housing material selection 100%
•
Selection of thermal coating 100%
•
Implement isogrid patterns 100%
•
Familiarization of software environment for analysis.
•
ProE 100%
•
Ansys 100%
•
Import methods 100%
38
Task Breakdown, cont’d.
•
Thermal analysis.
•
Receive determined component locations 100% Complete
•
Obtain relevant thermal constants 100%
•
Obtain relevant material properties 100%
•
Orbit propagation code for case determination 100%
•
Determine boundary conditions 100%
•
Generate thermal model for component heat sources 100%
•
Run simulations/verify results 50%
39
Supporting slides for mike 1
40
Satellite Structure
GPS Direct
Signal
Antennas
Solar
Panels
Lightband
Interface
High Gain Antenna
41
Supporting slides for
mike/dave 2
42
Project Scope
– Thermal
• Provide thermal models of Goldeneye with nodes for
each of the temperature critical components onboard
• Provide complete list of heat sources and their profiles
• Determine orbit hot and cold cases
• For each component and at each node of the thermal
models determine:
– Operating temperature: Temperature at which
the component will function and meet all
requirements
– Non-operating temperature: Component
specifications are not required to be met.
Component can be exposed in a power off
mode. If turned to power on mode, damage must
not occur
– Survival temperature: Permanent damage to the
component
– Safety temperature : Potential for catastrophic
damage
43
Thermal Analysis: Boundary Conditions
Boundary Conditions:
– Internal Heat Generation
•
•
•
•
IMU – 9.7 Watts (operational)
Computer – 9 -19 Watts
Battery < 1 Watt
Component Box 1 (ADNCS Microprocessor, Converter):
– Cold: 1 Watt
– Hot: 14 Watts
• Component Box 2 (Radios)
– Cold: 3 Watts
– Hot: 26 Watts
44
Thermal Analysis:
Future work
ANSYS
– Model is much to robust for computing resources
– Need to simplify our analysis
• Reduce node refinement at non-critical points
• Eliminate re-radiation between some internal components:
– Most likely from boxes to other boxes
• Shorten time steps (length of analysis)
– Currently doing 6 orbits
– Analyze Thermal Results
• Design changes if necessary
– Test Convergence / Accuracy
45
Top Level Requirements
Provide thermal histories for all temperature critical components under
hot and cold worst cases.
Requirement
Number
Requirement
Type
Verification Document
Status
THRM-1
Assemble list of material properties, temperature critical
component profiles
research
GEN-ANA-0001_A
verified
THRM-2
Provide thermal models of Goldeneye with nodes for
each of the temperature critical components
onboard
analysis
GEN-ANA-0001_A
verified
THRM-3
Determine hot case and cold case thermal boundary
conditions
analysis
GEN-ANA-0001_A
verified
THRM-4
Determine temperature history for each temperature
critical component
analysis
GEN-ANA-0001_A
verified
Thermal Boundary Conditions -Heat Fluxes
-Fluxes are date/time,
attitude, orbit dependent 
use Simulink/M-files
-Double quadruple integrals+ 832
lines=1.5 - 3 hrs run time per 1
orbit
47
Thermal Boundary Conditions - Albedo
Boundary Conditions:
-Fluxes are date/time, attitude, orbit
– Solar Flux: 1327 – 1414 Watts/m2
– Earth Albedo
– Earth IR
Source: http://www.tak2000.com/data/planets/earth.htm
Extracted from: Thermal Environments JPL D-8160
48
Hot case, Hot face
Hot case, Component Box (GPS receiver,
ADNCS, etc)
Allowable
Temperature Range:
-40 to 85 deg C
Hot case, Flight Computer
Allowable
Temperature Range:
-20 to 70 deg C
Cold case, Component Box: GPS, ADNCS, etc
Allowable Temperature Range: -40 to 95 deg C
52
Cold case, Flight Computer
Allowable Temperature Range: -55 to 125 deg C
53
Cold case, IMU
Allowable Temperature Range: -40 to 85 deg C
54