1. University of Minnesota Senior Design II Nanosat-V Final Design Review 6 May 2008 Minneapolis, MN 1

2. 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

3. Thermal Analysis (THRM) Subsystem Overview Thermal Analysis Team David Hauth Chuck Hisamoto Michael Legatt 3

4. 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

5. 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 Boxes Flight Computer

6. Thermal Analysis (THRM) David Hauth 6

7. 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

8. –Alodine Aluminum (6061 T6) Thermal conductivity:167 W/m 2 Specific Heat:896 J/kg-K Absorptivity/emissivity: Solar:.35 IR:0.1 –Emcore Triple Junction GaAs Solar Cells Annealed at 200 deg C Absorptivity/emissivity: Solar:.92 IR:.89 –Nusil CV10-2568 Controlled Volatility RTV Ablative Silicone Adhesive Operating Temperature Range (deg C): -115 to 240 8 Analysis Input: Material Properties

9. Internal Power Generating Components

10. Thermal Analysis Methodology 10

11. Thermal Analysis (THRM) Michael Legatt 11

12. 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

13. Beta Angle Solar Eclipse begins at Beta- star 13

14. Hot Case Cold Case Occurs at: -Beta=Beta-star -Lowest altitude=250km Occurs at: -Beta=0 -Highest altitude=1000 km 14

15. 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/m 2 15 Thermal Boundary Conditions

16. 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

17. Thermal Analysis (THRM) Chuck Hisamoto 17

18. 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 60N/A SA-60C GPS antennas-40 to 85-50 to 90 Sanyo N-4000DRL batteries0 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

19. Worst Hot case, Sun side Allowable Temperature Range: -115 to 240 deg C Cells Annealed at 200 deg C

20. Hot case, bottom Allowable Temperature Range: -115 to 240 deg C

21. Hot case, warmer near Standoffs Allowable Temperature Range: -115 to 240 deg C

22. Hot case, Isogrids, Standoffs Allowable Temperature Range: -115 to 240 deg C

23. Hot case, Battery Box Allowable Temperature Range: 0 to 40 deg C

24. Hot case, Component Box (Radio) Allowable Temperature Range: -20 to 60 deg C

25. Hot case, Inertial Measurement Unit Allowable Temperature Range: -30 to 60 deg C

26. 26 Thermal Performance – Hot Case Component Actual Temperature Range (deg C) Allowable Temperature Range (deg C) Pass/ Fail MinMaxMinMax Satellite Solar Panels/Cells-22.41122.76-115240Pass Cells Annealed:200 Battery Box21.58823.494040Pass Component Box28.95233.808-4085Pass (ADNCS, GPS, etc) Component Box50.74955.065-2060Pass (Radios) Flight Computer55.28158.769-2070Pass IMU44.64746.555-3060Pass

27. 27 Thermal Performance – Cold Case Allowable Temperature Range: -115 to 240 deg C

28. 28 Cold case, hot face/cold face Allowable Temperature Range: -115 to 240 deg C

29. 29 Cold case, Isogrids/standoffs Allowable Temperature Range: -115 to 240 deg C

30. 30 Cold case, Component Box (Radios) Allowable Temperature Range: -20 to 60 deg C

31. 31 Cold case, Battery Box Allowable Temperature Range: -30 to 60 deg C

32. 32 Thermal Performance – Cold Case Component Actual Temperature Range (deg C) Allowable Temperature Range (deg C) Pass/ Fail MinMaxMinMax Satellite Solar Panels/Cells-35.32929.645-115240Pass Cells Annealed:200 Battery Box-19.982-19.071-3050Pass Component Box-14.388-12.524-4085Pass (ADNCS, GPS, etc) Component Box-20.045-17.584-2060Fail (Radios) Flight Computer-14.949-14.442-5570Pass IMU-17.539-17.003-4085Pass

33. 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. 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. 35 Supporting Slides Follow

36. 36

37. 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. 2 nd ed. 2004. Kaminski, Deborah and Jensen, Michael. Introduction to Thermal and Fluids Engineering. 2005. Modest, Michael. Radiative Heat Transfer. 2 nd ed. 2003. 37

38. 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

39. 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

40. Supporting slides for mike 1 40

41. Satellite Structure GPS Direct Signal Antennas Solar Panels Lightband Interface High Gain Antenna 41

42. Supporting slides for mike/dave 2 42

43. 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

44. 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

45. 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

46. Top Level Requirements Requirement NumberRequirementTypeVerification DocumentStatus THRM-1 Assemble list of material properties, temperature critical component profilesresearchGEN-ANA-0001_Averified THRM-2 Provide thermal models of Goldeneye with nodes for each of the temperature critical components onboardanalysisGEN-ANA-0001_Averified THRM-3 Determine hot case and cold case thermal boundary conditionsanalysisGEN-ANA-0001_Averified THRM-4 Determine temperature history for each temperature critical componentanalysisGEN-ANA-0001_Averified Provide thermal histories for all temperature critical components under hot and cold worst cases.

47. 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

48. Boundary Conditions: –Solar Flux:1327 – 1414 Watts/m 2 –Earth Albedo –Earth IR Source: http://www.tak2000.com/data/planets/earth.htmhttp://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

49. Hot case, Hot face

50. Hot case, Component Box (GPS receiver, ADNCS, etc) Allowable Temperature Range: -40 to 85 deg C

51. Hot case, Flight Computer Allowable Temperature Range: -20 to 70 deg C

52. 52 Cold case, Component Box: GPS, ADNCS, etc Allowable Temperature Range: -40 to 95 deg C

53. 53 Cold case, Flight Computer Allowable Temperature Range: -55 to 125 deg C

54. 54 Cold case, IMU Allowable Temperature Range: -40 to 85 deg C