Upload
july
View
38
Download
0
Tags:
Embed Size (px)
DESCRIPTION
University of Minnesota Senior Design II Nanosat-V Final Design Review 6 May 2008 Minneapolis, MN. Project Objective. The aim of this project is to perform and validate thermal, structural and vibrational analyses on the Nanosat-5 satellite. - PowerPoint PPT Presentation
Citation preview
University of Minnesota
Senior Design II
Nanosat-VFinal Design Review
6 May 2008Minneapolis, 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 TeamDavid Hauth
Chuck HisamotoMichael 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
• 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– 2 for electrical components
• GPS Receiver, Radios, etc.
– 1 dedicated for batteries• Strict requirements for
coatings and narrower allowable temperature range
• IMU• Flight Computer
Battery Box
IMU
5
Component BoxesFlight Computer
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
– Alodine Aluminum (6061 T6)• Thermal conductivity: 167 W/m2
• Specific Heat: 896 J/kg-K• Absorptivity/emissivity:
Solar: .35IR: 0.1
– Emcore Triple Junction GaAs Solar Cells• Annealed at 200 deg C• Absorptivity/emissivity:
Solar: .92IR: .89
– Nusil CV10-2568 Controlled Volatility RTV Ablative Silicone Adhesive• Operating Temperature Range (deg C): -115 to 240
8
Analysis Input: Material Properties
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
• Use Beta angle
–Earth Albedo–Earth Infrared
12
Beta Angle
Solar Eclipse begins at Beta-star
13
Hot Case
Cold Case
Occurs at:
-Beta=Beta-star
-Lowest altitude=250km
Occurs at:
-Beta=0
-Highest altitude=1000 km
14
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
Thermal Boundary Conditions
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
ComponentOperating 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
Sanyo N-4000DRL batteries 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
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
26
Thermal Performance – Hot Case
ComponentActual Temperature Range
(deg C)Allowable Temperature Range
(deg C)Pass/Fail
Min Max Min Max Satellite Solar Panels/Cells -22.41 122.76 -115 240 Pass
Cells
Annealed: 200 Battery Box 21.588 23.494 0 40 Pass Component Box 28.952 33.808 -40 85 Pass
(ADNCS, GPS, etc) Component Box 50.749 55.065 -20 60 Pass
(Radios) Flight Computer 55.281 58.769 -20 70 Pass IMU 44.647 46.555 -30 60 Pass
27
Thermal Performance – Cold Case
Allowable Temperature Range: -115 to 240 deg C
28
Cold case, hot face/cold face
Allowable Temperature Range: -115 to 240 deg C
29
Cold case, Isogrids/standoffs
Allowable Temperature Range: -115 to 240 deg C
30
Cold case, Component Box (Radios)
Allowable Temperature Range: -20 to 60 deg C
31
Cold case, Battery Box
Allowable Temperature Range: -30 to 60 deg C
32
Thermal Performance – Cold Case
ComponentActual Temperature Range
(deg C)Allowable Temperature Range
(deg C)Pass/Fail
Min Max Min Max Satellite Solar Panels/Cells -35.329 29.645 -115 240 Pass
Cells
Annealed: 200 Battery Box -19.982 -19.071 -30 50 Pass Component Box -14.388 -12.524 -40 85 Pass
(ADNCS, GPS, etc) Component Box -20.045 -17.584 -20 60 Fail
(Radios) Flight Computer -14.949 -14.442 -55 70 Pass IMU -17.539 -17.003 -40 85 Pass
33
Design Conclusions
Hot Case
•All temperature critical components survive orbit within operating ranges•Heat accumulated on “hot side”
-Satellite slow spin maneuver-Addition/changes to coatings
Cold Case
•Radios component box is slightly out of storage temperature range.
-Need for heaters-Small generation needed
•All other components survive within range
34
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
35
Supporting Slides Follow
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
AntennasSolar
Panels
Lightband Interface
High Gain Antenna41
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
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: Boundary Conditions
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
Thermal Analysis: Future work
45
Top Level Requirements
Requirement Number Requirement Type Verification Document Status
THRM-1Assemble 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-3Determine hot case and cold case thermal boundary
conditions analysis GEN-ANA-0001_A verified
THRM-4Determine temperature history for each temperature
critical component analysis GEN-ANA-0001_A verified
Provide thermal histories for all temperature critical components under hot and cold worst cases.Provide thermal histories for all temperature critical components under hot and cold worst cases.
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
Boundary Conditions:– 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
Thermal Boundary Conditions - Albedo
-Fluxes are date/time, attitude, orbit dependent use Simulink/M-files -Fluxes are date/time, attitude, orbit dependent use Simulink/M-files
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
52
Cold case, Component Box: GPS, ADNCS, etc
Allowable Temperature Range: -40 to 95 deg C
53
Cold case, Flight Computer
Allowable Temperature Range: -55 to 125 deg C
54
Cold case, IMU
Allowable Temperature Range: -40 to 85 deg C