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Thermal, Telecommunication and Power Systems for a CubeSat

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1. 22 Figure 11 Thermal Vacuum Bakeout Profile 2 2004 sese 24 Figure 12 Possible Configuration of Pull Pin and Separation Switches adapted from 6 30 Figure 13 Ground track with ground stations seen enne 35 Lrpure T4 Access REPOT s i os aie a E E E pce a Mc Me 36 Fig re 135 Link Budget 5 eot ea ualet E a a 37 Figure 16 Access Report for the Four selected Ground Stations ees 38 Figure 17 Ground tracks of the CubeSat plotted using STK eeeeeeees 39 Figure 18 Ground track showing location of satellite connected to two ground stations Delft and Warsaw dere ccu a cto dE ppt D oe p Ca UO 41 Figure 21 Left 3D cube at time t 0 sec amp Right 3D cube at time t 172800 sec 44 Ergure 227 CubeSat Axes Ji a asete osos eee on ERU Ur OP IR DEAE iet eint md Pent OU d Nude 45 Figure 23 CubeSat sun Vector cdita oet ei gleaned a Queues 46 Figure 24 Aluminum Skeleton for Lab Option 2012 1 esee 47 Figure 25 Total expected daily access time as a function of launch date 56 Figure 26 Expected daily downlink access at each ground station as a function of launch date 57 Figure 27 Expected daily uplink access at each ground station as a function of launch date 57 Figure 28 Vacuum Chamber Model oe Ge de deus 61 Figure 29 Temperature
2. 13 SW 11 5 05 1 Magnetic Torquer 1 7 23 PDM 15 SW 13 5 05 1 Magnetic Torquer 3 7 23 PDM 17 SW 14 5 05 4 7 PDM 19 SW 15 8 3 BATT 0 5 7 PDM 21 SW 17 8 3 BATT 1 Payload 7 PDM 23 SW 20 opt 2 596 7 PDM 25 5V 5V Board Stack Connection 6 EPS Only 27 33V 3 3V Board Stack Connection 6 EPS Only Po 29 GND Magnetic torque 1 3 6 30 EPS CubeSat Sun Sensor 23 31 AGND Board Stack Connection 6 EPS Only 33 Batt POS DO NOT CONNECT 6 EPS 35 PCM IN Deployment Switch 1 amp 2 6 EPS input to PCM s and PDM switch on state Power Flow to Board Stack 37 Dummy Pull pin inserted off state 6 Load EPS 39 Unused Connection 6 Deployment switch 1 amp 2 off state 41 BCR Out BCR out to Deployment 6 EPS Switches 43 BCR Out BCR out to Deployment 6 EPS Switches 45 BATTERY Board Stack Connection 6 Only 47 RS422 RXA 7 49 RS422 RXB 7 51 OPTIONAL 7 V 2 SW 2 3 33 0 5 7 PDM 4 SW 4 3 33 1 7 PDM 115 6 SW 6 3 33 1 7 PDM 8 SW 7 3 33 4 7 PDM 10 SW 8 5 05 0 5 CubeSat Sun Sensor 7 30 PDM 12 SW 10 5 05 0 5 Gyroscope 7 61 PDM 14 SW 12 5 05 1 Magnetic Torquer 2 7 23 PDM 16 SW 14 5 05 4 GPS red 1 7 63 PDM 18 SW 14 5 05 4 GP
3. LAUNCH SAFETY REQUIREMENTS FOR AIR FORCE SPACE COMMAND ORGANIZATIONS 4 6 1 6 4 Flight Safety System Composition The entire FSS consists of the airborne and ground FTSs air borne and ground tracking systems and airborne and ground telemetry data transmission systems This publication only addresses the Launch Safety requirements for the ground components of the FSS Requirements for the airborne components of the FSS are in AFSPCMAN 91 710 The ground element of the FSS consists of the TDTS the RTS and the FTS The RTS includes a RSD system which displays vehicle flight performance data that the MFCO uses as the basis for a flight termination decision The TDTS provides for the transport of onboard launch vehicle position and performance data to the RSD and the Range User The CDS provides the MFCO the capability to terminate the vehicle s flight RTS Requirements An RTS is composed of the hardware software and manpower required to transmit receive process and display selected launch vehicle data This data when qualified as described below allows a MFCO to compare actual and nominal flight trajectories verify performance in conjunction with vehicle telemetry and identify violations of destruct criteria 89 6 5 6 7 1 6 7 2 6 7 3 6 7 3 1 6 7 3 2 6 7 3 3 6 7 3 4 6 7 3 5 6 7 4 6 7 5 6 7 5 1 6 7 5 2 6 7 6 With this information the MFCO knows when a flight rule has be
4. 0 59164 0 32546 40 00 0 0 22516 0 91848 0 3251 15 00 0 0 75032 0 5754 0 32547 45 00 0 0 24513 0 91335 0 32513 20 00 0 0 7627 0 55888 0 32549 50 00 0 0 26498 0 90778 0 32515 25 00 0 0 77471 0 5421 0 32551 55 00 0 0 2847 0 90178 0 32516 30 00 0 0 78635 0 52506 0 32553 00 00 0 0 30429 0 89536 0 32517 35 00 0 0 79761 0 50777 0 32556 05 00 0 0 32373 0 88851 0 32518 40 00 0 0 80848 0 49025 0 32559 10 00 0 0 34302 0 88124 0 32518 45 00 0 0 81897 0 4725 0 32562 15 00 0 0 36216 0 87355 0 32518 50 00 0 0 82907 0 45453 0 32564 20 00 0 0 38112 0 86545 0 32518 55 00 0 0 83878 0 43634 0 32566 25 00 0 0 39991 0 85693 0 32518 00 00 0 0 84808 0 41795 0 32568 30 00 0 0 41852 0 848 0 32519 95 Appendix 4 Lab Option Hardware List The following are images and details for the equipment purchased for the vacuum chamber experiment Halogen Lamp amp Bulb 35 Workforce 250 Watt Halogen Portable Work Light Home Depot Model 778 980 Home Depot Store SKU 778980 8 99 Aluminum Sheet 36 Alloy 3003 Aluminum Sheet Dimensions Thickness 0 0252 in Width 12 in Length 12 in MSC Industrial Supply Co MSC 09426057 4 52 PCB 37 PCBs Circuit Boards Single Sided Copper Clad Dimensions Thickness 1 32 in Width 6 in Length
5. 1 For analysis and testing they chose the simple aluminum structure that the Mechanical and 46 Structural Subgroup built which is shown in Figure 22 They believed that using this structure would be much easier to perform thermal analyses and form meshes during COMSOL analyses Using what was suggested for future groups a model was created to be simulated in COMSOL Figure 22 Aluminum Skeleton for Lab Option 2012 1 Before creating a model for analysis a layout had to be developed for the Lab Option In the chamber the CubeSat skeleton would be suspended from the aluminum stand created by the Mechanical and Structural Subsystem The stand and the skeleton would be positioned in front of the cryopump A halogen lamp will simulate the heat flux from the sun and be positioned on the other side of the cryopump In choosing which heat transfer module to use the best option was Heat Transfer with Surface to Surface Radiation Two other options had Radiation in Participating Media which implies that there is air or fluid in the system When the cryopump pumps air out of the chamber 47 it reaches temperatures close to absolute zero and pressures below 50 mTorrs Based on this knowledge the ambient temperature inside the chamber was set to 30 degrees Kelvin During testing five thermocouples measure the temperature at arbitrary points on the aluminum CubeSat skeleton Since the size and weight of the thermocouple
6. 70 5 2 2 WPI Ground Station Proposal Through research into the requirements from Cal Poly and the other requirement documents as well as the ground stations in the GENSO network a finalized list of hardware was created for both on board the CubeSat and the ground station at WPI A monetary budget of the hardware part number and price for all telecommunication related hardware has been created maximum and minimum cost analyses tables can be found in Appendix 5 The maximum estimated cost for implementation of a WPI Ground Station is 11 646 The antenna currently located on top of Atwater Kent has been approved to be relocated to the top of Higgins Laboratories to be used for the WPI ground station Facilities will need to be contacted to begin the bidding for a company to relocate the antenna from Atwater Kent to Higgins Laboratory The room currently reserved for the WPI ground station has been cleared out and is ready for the necessary equipment With a ground station at WPI the unique address of the frequency may be added to the GENSO network Every ground station has its own frequency for which an application with the Federal Communications Commission FCC needs to be completed Different current and future CubeSat mission orbits will be reviewed to see if other satellites will be able to use the WPI ground station for their telecommunication architecture This analysis will likely that the WPI ground station will not only be useful to t
7. Accessed 2013 54 Texas Towers Yaesu G 5500 TexasTowers com Online Available http www texastowers com g5500 htm Accessed 2013 55 Rakuten Linksys EtherFast BEFSR41 Broadband Router 1 x 10 100Base TX WAN 4 x 10 100Base TX LAN BEFSR41 LA Buy com Inc 2013 Online Available http www rakuten com prod linksys etherfast befsr4 broadband router 1 x 10 100base tx wan 4 x 10346262 html Accessed 2013 56 Martin RF Supply Bird Model 43P Thruline RF Wattmeter Peak Avg Kit New Online Available http www chuckmartin com category Bird 43 Wattmeters 2 Accessed 2013 57 Meters amp Test Equipment Universal Radio Inc 2012 Online Available http www universal radio com catalog meters 1739 html Accessed 2013 58 Ettus Research Quick Order National Instruments 2013 Online Available 78 https www ettus com product quick order Accessed 2013 59 Texas Towers Polyphaser IS 50NXC2 TexasTowers com Online Available http www texastowers com is5O0nxc2 htm Accessed 2013 60 DXengineering PolyPhaser Rotator Control Line Protectors IS RCT Auto Sales Incorperated 2013 Online Available http www dxengineering com parts ppr is rct Accessed 2013 61 Analog Devices Evaluation Board User Guide UG 245 17 March 2011 Online Available http www analog com static imported files user_guides UG 245 pdf Accessed 20 February 2013 62
8. P POD An overview of the power subsystem is shown by the flowchart in Figure 1 Management Distribution Users Generation Solar Array with integrated EPS Electronic Power System PDM Power OBC Distribution Module temperature Monitors solar panel output and sermons temp Adjusts power point to ADCS max mize output On off switching Over current protection Manages output andinput to Telecom battery Optimizes voltage Multiple bus distribution Battery protection gt Instrument Sensors Battery interface Board a Battery Cell Battery Cell Figure 1 Power Subsystem Solar Array Power generation duties on CubeSats are covered by solar arrays due to the abundant solar energy available in orbit there is over four times the energy available to satellites compared to ground based panels As such the panels used on CubeSats are small and light while still providing plenty of power to users Many solar panels designed for operation in space use gallium arsenide cells because of its higher efficiency compared to silicon and a lightweight substrate made of fiberglass Printed Circuit Board PCB aluminum carbon fiber or an alternative composite The power level generated by the cells is a function of their efficiency area cell density and temperature While the physical qualities of the panels are only affected by degradation over time temperature of the pane
9. a detailed procedure will need to be prepared for each of these 1 Mount lamp and CubeSat model so that they are thermally isolated from the chamber walls using monofilament fishing line 2 Place thermocouples in selected locations to measure the thermal profile of the CubeSat 3 Run vacuum chamber with lamp on and take data both from the sides exposed to the light and not exposed to the light The results of this experiment can be compared to the results of the COMSOL simulation of the same situation 3 3 5 Thermal Control System The thermal control system ensures that the operational temperatures of each component are not exceeded The following table shows the operating temperatures for the equipment on board The data has been collected from the specifications given by the manufacturers Component Operational Temp C Reference OBC 40 85 19 ADC board No Data 20 Course Sun Sensor 3 or 5 40 100 21 Fine Sun Sensor 25 50 9 Gyro 40 105 22 Magnetic Torquers 3 No Data 23 Magnetometer 30 85 24 GPS 0 50 25 SphinX NG Instrument Self regulating thermal abilities EPS Board 40 85 6 PDM Board 40 85 7 Battery 10 50 26 Front Solar Panels 40 80 27 Side Solar Panels 40 80 28 UHF Transceiver 30 70 10 Table 5 Operational Temperatures 49 From Table 5 it is evident that certain components have more limited r
10. characteristics Dipole Length 0 55 m Length Wave Length 0 799 Efficiency 5596 Refraction Model ITU R P 834 4 Range limit 3000 km Each of the ground stations incorporated a medium complexity receiver model with the following constraints Gain 32 8 dB Line Loss 4dB Antenna noise from sun atmosphere rain computed by STK LNA low noise amplifier noise figure 1 2 dB Refraction Model ITU R P 834 4 Min Elevation angle 5 deg Doppler Shift 25 kHz 7 ITU R P 834 4 Effects of tropospheric refraction on radiowave propagation International Telecommunications Union 67 40 Reports From the aforementioned modeling constraints STK can calculate the maximum access and link conditions for the model This data is recorded within STK and can be exported in a number of report formats The reports providing the needed information to accurately assess the telecommunication system capability The Access Reports consist of the duration and occurance of links between the satellite transmitter and ground station receivers This is affected by the orbital parameters and the field of view of the satellite and ground station especially the effective range of the transmitter and the minimum elevation of the field of view from the ground station Figure 18 shows the CubeSat within range of the Delft and Warsaw ground stations STK calculates the coverage for both although when considering the amount of data received it
11. forward link is the link from a fixed location such as a base station to a mobile user A forward link will contain an uplink base station to satellite and downlink satellite to mobile user if a communications relay satellite is involved A reverse link is the link from a mobile user to a fixed base station In the case of a communications relay satellite the reverse link will have both an uplink mobile station to satellite and a downlink satellite to base station 12 A crosslink is the link between a satellite and another satellite These links are part of a larger network within the ground station a telecommunications network This network is created through a collection of terminals links and nodes that all connect to ensure telecommunication between terminals and ground stations A unique address is created for each terminal in the network so messages will be sent correctly Address space is the collection of these unique addresses in each network 18 The geometry of the links and the ground station create the telecommunication architecture There are different types of architectures each with their own advantages and disadvantages Store and forward is an architecture for relaying communications by satellite For this the satellite orbits at a low altitude and receives data that is stored in its memory The data is transmitted when the satellite is in view of a receiver ground station The advantages of this architecture are
12. http www mouser com ProductDetail MG Chemicals 588 qs sGAEpiMZZMuwe2fyn7eX6 252bV QtT7NeG 252bLat9KTJDReH E 3d Accessed 2013 38 OMEGA Ready Made Insulated Thermocouples with Kapton PFA Glass Braid Insulation and Molded Connectors Omega Engineering Inc 2013 Online Available http www omega com ppt pptsc asp ref 5LSC_SSRTC Accessed 2013 39 McMaster Carr Thermocouple and RTD Wire amp Converters Online Available http www mcmaster com fstcatalog 119 601 21q0mjw Accessed 2013 40 McMaster Carr Thermocouple and RTD Probes amp Connectors Online Available http www mcmaster com fstcatalog 119 600 2lqlzwl Accessed 2013 41 K J Lesker Power Feedthroughs Weldable 500 Volts Kurt J Lesker Company 2013 Online Available http www lesker com newweb feedthroughs power feedthroughs cfm pgid 500v weld Accessed 2013 42 T S Tuli N G Orr and D R E Zee Low Cost Ground Station Design for Nanosatellite Missions University of Toronto Institute For Aerospace Studies 2006 Online Available http www utias sfl net docs canx2 amsat 2006 pdf Accessed 2013 43 CubeSatShop com Small Ground Station ISIS 2013 Online Available http www cubesatshop com index php page shop product details amp flypage flypage tpl amp product_id 24 amp category_id 3 amp option com_virtuemart amp Itemid 72 amp vmechk 1 amp Itemid 72 Accessed 2013 44 The Department of
13. inside the CubeSat there is a thermal control system to regulate the temperature so that it does not depend purely on its orientation There are two categories of thermal control systems active and passive Active systems have the capacity to turn on or off to adjust the temperature in a more precise manner whereas 25 passive systems are fixed in place and heat is transferred through a natural process rather than a mechanical one Examples of active systems are heat pumps louvers or electric heaters or coolers which are commonly used in larger spacecraft such as those transporting humans or a very sensitive payload 15 Passive systems can be in the form of heat pipes thermal coatings blankets or radiators used in many different spacecraft especially in smaller simpler spacecraft such as a CubeSat or other small satellites 15 Passive thermal control will be used for this CubeSat Based on analysis of the model of the CubeSat performed in COMSOL the use of thermal coatings and a well placed radiator will suffice for the majority of the life of the CubeSat The P POD requirements state very specific temperature ranges that the CubeSat must survive in In order to meet these requirements the most careful analysis and testing will need to be focused on this portion of the mission 26 3 Methodology 3 1 Power Subsystem Distribution The electrical boards associated with the CubeSat can be stacked vertically and interfac
14. 0033 CS Side Solar SP L S2U 0031 28 Panels ClydeSpace CS UHF 10 ISIS Transceiver Transceiver Communicati on UHE ISIS Antenna Transceiver 49 IC910 ICOM Radio Ground MercuryRotor YASU Station driver K9612 the two ports 50 1200 amp Kantronics 9600 51 MixW TNC K8055 52 Velleman Velleman SP7000 SSB 53 ELECTRONI CS 105 Yaesu G 5500 54 AZ EL Yaesu Controller Ground Station Control Control Computer Computer External Hard Terabyte Drive Laptop SDR Laris Control Wireless 5 DEESA Network Router Uninterruptab le 1000AVR Power supply 1692 12 Volt Power supply 1460 1PS 12 Volt Power supply RF Directional 95 57 Thruline Watt Meter Mdl 43 58 USRP B100 58 USRP 1 SDR 106 58 WBX USRP daughter board 58 TVRX USRP daughter board 2 Coax 59 protectors Lightning Polyphaser suppressors IS 50NX C2 2 Cable 60 protectors Lightning Polyphaser suppressors IS RCT 107 Appendix 8 CubeSat Bus Wiring Diagrams This appendix provides a labeling of all the connection points between the CubeSat bus and the devices attached to the CubeSat Below are the figures that show the locations of the headers the pins of the headers and the labeling of the wires or connection points of th
15. 31 31 33 39 42 42 44 46 48 49 51 51 51 54 55 55 59 60 60 65 69 69 69 69 71 71 71 72 73 74 Appendices Appendix 1 Requirements per Reference Documents Appendix 2 Definition of Key Terms Appendix 3 Normalized Beta Values over 24 Hours at 5 minute Step Appendix 4 Lab Option Hardware List Appendix 5 WPI Ground Station Appendix 6 Other Ground Station Hardware Appendix 7 Complete Hardware Parts List with Images Appendix 8 CubeSat Bus Wiring Diagrams Appendix 9 CubeSat Bus Power Board Pin Assignments Appendix 10 On Board Computer OBC 64 80 80 92 95 96 98 100 103 108 112 120 vi Table of Figures Pigure L Power Subsystems ienai e ao d Ea a eru abite e NE cic de added 5 Figure 2 Solar panel output as a function of temperature 5 eee 6 Figure 3 Clyde Space 3U EPS 6 iis aec ote deseo fo rtu platte tonne de pte tco copa ois 7 Figure Clyde Space PDM 7 awe E eo E e SL s 8 Figure 5 Location of Deployment Switches 2 201 D einen en enne eene 10 Figure 6 Location of Access Port for 3U CubeSat 2 2011 esee 11 Figure 7 Telecommunications suDSystelm ee eee e resi t ce e e dete eerta te e Pes nels 14 Figure 5 ISIS deployable antenna 9 uis otro itor bte ridere eps tecto eu oreste eee eons 16 Figure 9 ISIS UHE VHF transceiver 10 ecce eee Retreat dae asa thee itera 17 Figure 10 P POD and CubeSat Environment Tests Thermal 2011 3
16. 7 23 pin 29 Coarse Sun Red Analog AI ADC board pin and 21 Sensor 1 Input connector TBD Black Ground CubeSat bus header 2 7 21 pin 30 Coarse Sun Red Analog AI ADC board pin and 21 Sensor 2 Input connector TBD Black Ground CubeSat bus header 2 7 21 pin 30 Coarse Sun Red Analog AI ADC board pin and 21 Sensor 3 Input connector TBD Black Ground CubeSat bus header 2 7 21 pin 30 Coarse Sun Red Analog AI ADC board pin and 21 Sensor 4 Input connector TBD Black Ground CubeSat bus header 2 7 21 pin 30 Solar Array 1 1 Power EPS Board SA1 6 Y connector pin 1 2 Ground EPS Board SA1 6 connector pin 2 3 Temperatur EPS Board SA1 6 e connector pin 3 117 Telemetry Solar Array 1 Power EPS Board SA1 6 2 Y connector pin 4 2 Ground EPS Board SA1 6 connector pin 5 3 Temperatur EPS Board SA1 6 e connector pin 6 Telemetry Solar Array 1 Power EPS Board SA2 6 3 X connector pin 2 Ground EPS Board SA2 6 connector pin 2 3 Temperatur EPS Board SA2 6 e connector pin 3 Telemetry Magnetometer 1 CubeSat bus header 2 7 62 GND pin 30 2 CubeSat bus header 2 7 62 33V pin 1 3 I2C ADC board pin and 62 I2C DATA connector TBD 4 Dc RC ADC board pin and 62 CLOCK connector TBD Gyroscope P1 PDD 5 05V CubeSat bus header 2
17. 7 61 pin 12 P1 GND Ground CubeSat bus header 2 7 61 pin 32 P1 MISO Master In SPI ADC board pin and 61 Slave Out connector TBD P1 GND Ground CubeSat bus header 2 7 61 pin 32 P1 CS Chip Select SPI ADC board pin and 61 connector TBD P2 GND Ground CubeSat bus header 2 7 61 pin 32 P2 CLK SPI SPI ADC board pin and 61 CLOCK connector TBD P2 GND Ground CubeSat bus header 2 7 61 pin 32 P2 MOSI Master Out SPI ADC board pin and 61 Slave In connector TBD P2 GND Ground CubeSat bus header 2 7 61 pin 32 GPS Green 1 TBD UART ADC board pin and 63 connector TBD Green 2 TBD UART ADC board pin and 63 118 connector TBD Green 3 TBD UART ADC board pin and 63 connector TBD Green 4 TBD UART ADC board pin and 63 connector TBD Green 5 TBD UART ADC board pin and 63 connector TBD Green 6 TBD UART ADC board pin and 63 connector TBD Green 7 TBD UART ADC board pin and 63 connector TBD Black Ground UART ADC board pin and 63 connector TBD Red 1 5 05V CubeSat bus header 2 7 63 pin 16 Red 2 5 05V CubeSat bus header 2 7 63 pin 18 CubeSat Sun TBD Analog AI ADC board pin and 30 Sensor Input connector TBD TBD Analog AI ADC board pin and 30 Input connector TBD TBD Analog AI ADC board pin and 30 Input connector TBD TBD Analog AI ADC board pin and 30 Input connector TBD TBD 5 05V CubeSat bus he
18. Aerospace Engineering and Engineering Mechanics FASTRAC Mission Plan The University of Texas at Austin 2004 Online Available http fastrac ae utexas edu documentation FASTRAC_missionplan_FCR pdf Accessed 2013 45 PolySat Ground Station Mantra amp Word Press 2012 Online Available 77 http polysat calpoly edu the lab ground station Accessed 2013 46 H W Oelze Magnetic Torquers for Micro Satellites ZARM Technik AG Bremen Germany 47 F Chip USB to RS422 Serial Converter Cable Datasheet FT 000116 Future Technology Devices International Limited FTDI Glasgow UK 2012 48 FTDI UM232R USB Serial UART Development Module Datasheet FT 000051 Future Technology Devices International Ltd 2002 49 VHF UHF AII Mode Transceiver IC 910H ICOM Japan 2005 50 Kantronics KPC 9612 Packet Communicator Kantronics Inc 2011 Online Available http www kantronics com products kpc9612 html Accessed 2013 51 B Tracey KD5TFD Station SDR 1000 Software Defined Radio 8 March 2005 Online Available http www ewjt com kd5tfd station sdr1000 html Accessed 2013 52 Velleman EXTENDED USB INTERFACE BOARD Velleman 2013 Online Available http www velleman eu products view country2gb amp lang en amp id 364910 Accessed 2013 53 eHam net eHam net Classifieds Detail eHam net LLC 2011 Online Available http www eham net classifieds detail 379484
19. Flux Density The transmitter s radiated power over the surface area of a sphere whose radius equals the distance between the transmitter and receiver FSK Frequency Shift Keying Modulation technique in which digital information is carried by discrete changes in frequency of a carrier wave g T Ratio of gain over temperature GENSO Global Educational Network for Satellite Operations GPS Global Positioning System Ground Tracks projection of the satellite s orbit onto the surface of the Earth H1 Header 1 PC Inter Integrated Circuit Consists of a data line and a clock line allowing for a connection of a low speed device to a motherboard or an embedded system ICD Interface Control Document LNA Low Noise Amplifier Low gain amplifier optimized to minimize the amount of noise the amplifier adds to the signal LV Launch Vehicle Mbps Mega Bytes per Second MPE Maximum Predicted Environment MPPT Maximum Power Point Tracking MQP Major Qualifying Project MUA Material Usage Agreement NASA National Aeronautical Space Association OBC On Board Computer PCB Printed Circuit Board PCM Power Conditioning Modules PDM Power Distribution Module PMAD Power Management and Distribution P POD Poly Picosatellite Orbital Deployer 93 QPSK Quadrature Phase Shift Keying Form of Phase Modulation in which the carrier phase can be one of 4 possible phases Has the same Bit error rate as BPSK but for 1 2 the bandwidth R A
20. Son Luis Obispo CA 93407 MUR PEMEHONSIDIESIGNED AY R a MUN PART NAME TOLERANCES ORAWN BY PLACEMENT OPTIONS FOR DEPLOYMENT T SWITCHES AND SEPARATION SPRINGS ANGULAR 20 5 A APPROVED BY E faces AND CUBESAT SPECIFICATION CORNERS DRAWING NOT TO SCALE DATE 0277 11710 SHEET J OF 1 Figure 5 Location of Deployment Switches 2 2011 10 s18938uutiur ur ase suorsuauiip my ip Sjegaqn e uo payeiodioour aq ysnw youms yuawAojdap euo 3sea yy GOd d 9n jo apisur ayy uonoj peus syre4 au uey 13430 s3uesuoduioo jeuis3xe ON 7 jegaqn au jo 133u82 2uauloa8 au ut pa3e20 uia3s s 83euipaooo 3egaqn SSLON TVNOILIGGV Z HO4 l vi3Q 12v1NOO d4JQ0NY1S wv e130 SU LYSIE Figure 6 Location of Access Port for 3U CubeSat 2 2011 11 The CubeSat shall also be self contained providing its own power sequencing and wiring 3 The requirement of the CubeSat being able to provide its own sequencing refers to it being able to carry out its tasks without constant commands from a ground station The electrical wiring of the satellite specifically its insulation shall also be inspected for flammability prior to its launch 3 If the wire insulation is chemically and physically similar to a material found to be acceptable by NASA STD 1 6001 then the material may be used without testing and justified on an approved Material Usage Agreement MUA After the thermal vacuum bakeout test is
21. Sparkfun Triple Axis Magnetometer Breakout HMC5883L 3 March 2013 Online Available https www sparkfun com products 10530 Accessed 3 March 2013 63 Surrey Satellite Technology LTD Space GPS Receiver SGR 05U 9 November 2012 Online Available http www sstl co uk getattachment 97ae8ccc 024d 4376 a99d 7d3c2266a7f7 SGR O5U 05P Accessed 20 February 2013 64 Tyvak Intrepid Pico Class CubeSat System Board Online Available http www tyvak com products Pico IntrepidS ystemBoard 2012July pdf Accessed 20 February 2013 65 R Simmons MOLA Analysis Definitions 20 December 1995 Online Available http analyst gsfc nasa gov ryan MOLA definit html Accessed 21 February 2013 66 CubeSat Wikipedia 9 April 2013 Online Available http en wikipedia org wiki CubeSat Accessed April 2013 67 I R Assembly Recommendation ITU R P 834 4 Effects of tropospheric refraction on radiowave propagation International Telecommunication Union 2003 79 Appendices Appendix 1 Requirements per Reference Documents The following eight lists are each a governing requirements document provided by institutions that help regulate what can and cannot be flown on a CubeSat Each has been thoroughly read through and streamlined to determine what is directly related to the SphinX CubeSat mission CubeSat Design Specification Rev 12 The CubeSat Program Cal Poly SLO 2 2 4 1 FCC requires all CubeSats w
22. dipole turnstile and patch antennas The antenna system best suited to the mission is the ISIS deployable UHF VHF antenna system This allows the use of one of three configurations four monopole antennas two dipole antennas or one turnstile For this mission the dipole configuration was chosen because of its ability to run both a UHF transmitter and a VHF receiver from the same unit providing a compact solution Along with the semi omnidirectional profile of the dipoles which will limit the pointing losses regardless of orientation this antenna offers a comprehensive solution to the antenna choice for the CubeSat S Band Secondary Downlink To further increase the downlink data rate a secondary radio operating on the 2 4 GHz S band is being considered The Clyde Space STX is a transmitter designed specifically for this purpose on CubeSats and could significantly enhance the transmission capabilities of the CubeSat The S band allows for data rates up to 2 Mbps however it comes at the price of reliability power and cost There are also fewer ground station options for S band CubeSats although the band is widely used on larger commercial satellites Modulation is handled by QPSK Quadrature Phase Shift Keying which is partly responsible for the high data rate QPSK uses four phases of modulation as opposed to two for BPSK so it can handle twice the data for the same bandwidth and Bit Error Rate BER The downside is the complexity
23. during the mission are put in place to prevent radio interference between the LV and other satellites A frequency application must be filled out approved and a license issued for a specific frequency to be used for communication between the satellite and ground station It has become necessary to have requirements that ensure redundancy in place so that in the event of error or failure there are back up commands in place Another command prerequisite is to make sure the CubeSat has the capability for Global Positioning System GPS tracking and real time on board system updates to assure the hardware is working correctly Requirements are in place to assure that all Commercial Off the Shelf COTS hardware used is in accordance with all the NASA and Cal Poly standards There are LV requirements to assure the CubeSat is installed correctly onto the LV This means that the CubeSat must meet vibration testing limits and assures that the satellite does not interfere physically or electronically with the mission of the LV These requirements also demand any testing that will assure the CubeSat s ejection from the LV goes smoothly These tests include but are not limited to testing the CubeSat switch to ensure the power stays off and that the ejection portion of the mission plan will not interfere with the LV s mission or the LV physically 2 2 2 Hardware The purpose of the telecommunications system in a CubeSat is to provide a link between the OBC
24. f Mea W W W OBC OBC Tyvak Intrepid 0 300 0 30 0 20 ADC board Clyde Space ae ae 0 100 0 10 0 10 Course Sun ComTech CubeSat Sun Sensor 5 AeroAstro Sensor 0 000 9 00 0 00 295 SSBV Space p and Ground Fine Sun Sensor 0 140 0 14 0 14 7 5 26 oF Sensor S 5 5 50 ystems ADC Gyro Surrey ADXRS450 0 030 0 03 0 03 6 0 Ts Magnetic Zarm MTO 5 1 Torquers 3 Technik AG Optimized 0 213 0 2173 UA 33 0 Magnetometer Honeywell HMC5883L 0 000 0 000 0 00 0 100 2 16 3 6 Surrey Satellite GPS SGR 05U 0 800 0 80 0 80 16 0 5 Technology USLLC Payload Instrument Sphinx NG XRAY 8 000 8 00 1 00 TBD TBD ps 3 3 5 current EPS Board Clyde Space CS 3UEPS2 NB 0 100 0 1 0 1 raw protectio i battery 24 switches CN SWT 0035 each 3 3 512 Boon PDM Board Clyde Space CS 0 160 0 16 0 16 with own battery current telemetry Battery Clyde Space CS SBAT2 30 1 25 Ah 8 2 Front Solar SP L F2U 12 26 Panels Clyde Space 0033 CS 5 200 3 35 0 752 Side Solar SP L S2U 12 26 Panels Clyde Space 0031 CS 15 600 6 70 0 00 7 53 Telecom JHE ISIS Transceiverand 2100 0 20 0 00 TBD 6 5 12 5 Transceiver Antenna Total 12 005 9 705 2 370 Total Available 50 800 40 050 30 000 Margin 38 795 30 345 27 630 Table 6 Updated Power Budget 52 According to established CubeSat requirements all power must be off for launch Once launched and ejected from the launch vehicle the battery must power everything until de tumbling is compl
25. however the user specified voltage must be applied to pins H2 51 and H2 52 of the CubeSat bus Each switch also has a recommended current trip of 0 25 Amps 0 5 Amps 1 Amp or 4 Amps The current trip is a maximum current that the switch can handle before it trips a circuit breaker A complete list of each of the switches maximum voltage and current trip are located in Table 9 and Table 10 of Appendix 9 Three of the switches 7 14 and 19 also differ from the other 21 switches as they can connect three devices to one switch The power return for all switches must also go to the ground pins H2 29 H2 30 or H2 32 of the CubeSat bus The boards in the stack and the OBC receive power through pins 35 and 36 of the second header of the CubeSat bus and as a result are not powered on or off through a switch on the PDM board Once the pull pin is removed and the deployment switches are no longer actuated the 28 boards in the stack receive power without interruption 6 The only components that require switches to power them on or off are the three magnetic torquers the magnetometer the GPS the payload and the gyroscope 3 1 1 Telemetry Communication Each SA connector on the EPS board receives sensor information about the temperature voltage and current of each solar panel and sends it through the CubeSat bus The sensor information for the 3 3 Volt bus the 5 Volt bus and the 8 3 Volt battery bus are also sent to the CubeSat bus along with the g
26. of QPSK and the higher power needed STX S Band Transmitter from ClydeSpace Mbps Mega Bytes per Second 32 Antenna options for S band are primarily limited to patch antennas due to the higher power required for equivalent bandwidth compared to UHF or VHF Although patch antennas are simpler and easier to implement than deployable dipoles or similar they are highly directional producing a beamwidth of approximately 65 with significant losses outside of this range This proves problematic for this mission proposal since the pointing requirements for the Cubesat are sun dependent and there is no primarily nadir Earth facing side to mount the antenna A possible solution is the use of ground stations near the equator where the Z side of the satellite would be within the pointing requirements 3 2 2 Ground Station Network Ground tracks are the projected path of a satellite s orbit on the surface of the Earth which traces the movement of an imaginary line between the satellite and the center of the Earth Also it is considered a set of points that the satellite will pass directly over in the frame of reference of a ground observer 16 There are a few parameters that can cause variations in the ground tracks such as orbital period and orbital inclination A satellite with an orbital period of an integer fraction of a day 24 hours 12 hours 8 hours etc will for the most part follow the same path day to day The ground tracks w
27. should be noted that data will be redundant when two ground stations are connected to the Cubesat simultaneously However this does have the benefit of being able to check the stream of data using two or more sources for increased accuracy 41 Figure 15 is the access report data from a 24 hour period In this time the CubeSat would make approximately 15 orbits around the earth clearly demonstrating that there is marked room for improvement in coverage The link budget supplies vital data of the various link characteristics such as EIRP Effective Isotropically Radiated Power RIP Received Isotropic Power and BER The values are calculated by STK and can be analyzed to determine further constraints in the simulation and specify the proper components The output values can be seen in Figure 16 3 3 Thermal Subsystem 3 3 1 Thermal Analysis from External Sources CubeSat Attitude Determination and Control MQP Team 2011 2012 18 used STK to find spacecraft sun vectors for multiple orbits The spacecraft sun vector is coordinates of the satellite from its location to the sun These values were converted to beta values where the magnitude of the vectors equal 1 Once the time dependent beta values were found the next task was to import them into COMSOL to begin the next step in the thermal analysis With these values the solar flux can be made time dependent Since the position of the CubeSat is changing over the course of its orbit the so
28. that a low cost launch vehicle and low cost satellite can be used due to the low altitude and wider antenna beam width which reduces the satellite antenna size and stabilization requirement The disadvantage is there is a long message access time and transmission delay since they are waiting for the satellite to pass into view 13 Geostationary orbit is used by communication relay satellite systems and meteorological satellites The satellite is placed in a near zero degree inclination orbit at about 36 000 km altitude The orbit of the satellite is equal to the period of Earth s rotation which is one of its advantages The cost of a ground station is less because there is no need for antenna pointing control so it is a stationary network A stationary network is easier to set up monitor control and there is no need to switch satellites since the satellite is always in view The disadvantage is the lack of coverage above 70 degree latitude and the high launch cost There is also a delay time for propagation to and from the orbit which can cause problems 13 A Molniya orbit is used to cover the northern Polar Regions with the satellites in highly elliptical orbits The specification for the orbit is an apogee of 40 000 km which is over the North Pole a perigee of 500 km and an inclination angle of 63 4 degree The period of the orbit is 12 hours and since the orbit is highly elliptical the satellite spends about 8 hours of each period over
29. 3 40 00 0 0 45512 0 82892 0 3252 10 00 0 0 143436 0 93482 0 32485 45 00 0 0 47311 0 81878 0 32522 15 00 0 0 123022 0 93772 0 32488 50 00 0 0 49086 0 80825 0 32524 20 00 0 0 102553 0 94017 0 3249 55 00 0 0 50838 0 79734 0 32526 25 00 0 0 082036 0 94218 0 32491 00 00 0 0 52565 0 78605 0 32529 30 00 0 0 06148 0 94374 0 32492 05 00 0 0 54267 0 77439 0 32532 35 00 0 0 040893 0 94486 0 32492 10 00 0 0 55942 0 76236 0 32535 40 00 0 0 020282 0 94552 0 32493 15 00 0 0 5759 0 74998 0 32537 45 00 0 0 00034 0 94574 0 32493 20 00 0 0 59211 0 73724 0 32539 50 00 0 0 02097 0 94551 0 32493 25 00 0 0 60804 0 72415 0 32541 55 00 0 0 0416 0 94482 0 32493 30 00 0 0 62367 0 71072 0 32543 00 00 0 0 06221 0 94369 0 32493 35 00 0 0 63901 0 69696 0 32544 05 00 0 0 08279 0 9421 0 32494 40 00 0 0 65405 0 68286 0 32544 10 00 0 0 10333 0 94007 0 32495 45 00 0 0 66879 0 66844 0 32544 15 00 0 0 12383 0 93758 0 32497 50 00 0 0 6832 0 6537 0 32544 20 00 0 0 14426 0 93465 0 32499 55 00 0 0 6973 0 63864 0 32544 25 00 0 0 16462 0 93127 0 32502 00 00 0 0 71107 0 62327 0 32544 30 00 0 0 1849 0 92745 0 32505 05 00 0 0 7245 0 6076 0 32545 35 00 0 0 20508 0 92318 0 32507 10 00 0 0 73759
30. 3 793 17 Nov 2012 10 06 02 038 17 Nov 2012 10 06 02 038 16 Nov 2012 20 39 22 804 m Duration sec 563 325 629 405 852 786 447 201 790 327 793 144 447 201 852 786 679 365 4076 189 Duration sec 726 465 859 662 699 760 565 193 840 230 805 505 371 701 371 701 859 662 695 502 4868 516 m 30 Graphic BE 2D Graphic M Access Report Acc Al Satelite1 89 500 47 000 Figure 14 Access Report 16 Nov 2012 17 00 00 000 Time Step 60 00 sec Figure 13 is a screen shot of the ground tracks over the ground stations The thin light blue line is the ground track that the satellite is on The thicker different colored lines are the portion of the ground tracks during times the satellite is in contract with a ground station The different colors correspond with the different ground stations Figure 14 is a screen shot of the access report that STK compiles from the information on the satellite orbit and the ground stations From this information ground stations were streamlined to only focus on how much data could be transmitted The final choices are listed below in Table 3 Institution Location Country Nicolaus Copernicus Warsaw University of Warsaw Poland Astronomical Center Technology Delft Command Ground Delft University of Delft Netherlands Station DCGS Technology FASTRAC University of Texas at Austin Austin Texas Cal Poly Earth St
31. 5 Velleman Switch 40 SP7000 SSB ELECTRONICS amp 330 Ground Station Control Computer comp 800 Terabyte External Hard drive comp 70 2 Linux Laptops for SDR Control comp 302 Kantronics 3 plus Modem comp 30 Wireless Network Router BEFSR41 comp Free Uninterruptable Power Supply 1000AVR power 100 12 Volt Power supply 1692 power 345 RF Directional Thruline Watt Meter Mdl 43 power 155 USRP 2 SDR interface 850 USRP 1 SDR interface 700 USRP daughter board WBX interface w above bus USRP daughter board TVRX interface 200 2 Coax protectors Polyphaser IS 50NX C2 Lightning 62 suppressors lightning rod 2 Cable protectors Polyphaser IS RCT Lightning 130 suppressors lightning rod 7 663 99 Appendix 6 Other Ground Station Hardware The following tables are compilations of the hardware and software being used by each of the ground stations that have been used in calculating the ground tracks and data link budget The information was used to determine antenna information for the data information The lists were also used in determining hardware for the proposed WPI Ground Station By using similar or the same hardware and software the ground stations will more easily be able to communicate Nicolaus Copernicus Astronomical Center Warsaw University of Technology Warsaw Poland 42 hardware use two Yagi Uda antennas for UHF Antenna s
32. 9 in Distributor Mouser Electronics Mouser Part 590 588 7 46 Thermocouples 38 Ready Made Insulated Thermocouples with Kapton PFA Glass Braid Insulation and Molded Connectors Thermocouple type k insulate glass braid exposed junction AWG Gage 24 Diameter 0 020 in Miniature connector Omega Part Number 5SRTC GG K 24 36 50 00 5SRTC 1 5SC Extension Wires 39 24 Gage Solid Thermocouple Wire with FEP Insulation and Jacket K type M McMaster Carr Part Number 3870K32 1 20 ft Connectors for the thermocouples to the chamber feedthroughs 40 Panel Mount Thermocouple Female Connectors Panel Cutout 7 8 in diameter McMaster Carr Part Number 3869K48 10 82 Chamber electrical feedthrough 41 Power Feedthroughs Weldable 500 Volts Stainless Steel 4 pins Kurt J Lesker Co Part Number EFT0042031 81 50 97 Appendix 5 WPI Ground Station Below are two cost estimates of all required hardware excluding the antenna needed to create a WPI Ground Station The difference between Upper and Lower is simply pricing All part numbers are the same between the two Lower is the used version of the parts The decision to show two different cost estimates was to establish a range of possible totals for all hardware Upper values are all new parts from the manufacturer or commercial sources while Lower is used hardware from various d
33. A N Relative Angle of the Ascending Node RBF Remove Before Flight RF Radio Frequency RIP Received Isotropic Power The power of a signal at the receiving station relative to an isotropic antenna Computed by taking the EIRP and subtracting the free space loss atmospheric loss and pointing loss SA Solar Array S Band IEEE standard for radio waves frequencies that range from 2000 to 4000 MHz SHF Super High Frequency STK Systems Took Kit TBD To Be Determined TVB Thermal Vacuum Bakeout UHF Ultra High Frequency frequencies that range from 300 to 3000 MHz USAF United Stated Air Force USB Universal Serial Bus VHF Very High Frequency frequencies that range from 30 to 300 MHz WPI Worcester Polytechnic Institute Worcester Massachusetts 94 Appendix 3 Normalized Beta Values over 24 Hours at 5 minute Step The following tables show a sampling of the x y and z normalized coordinates for the vector between the spacecraft s center and the sun over time in five minute increments over the span of one day This vector can be seen in Figure 21 Electronic file available in the project file archive Time Time UTCO i y A UTCO T Y 00 00 0 0 184064 0 9277 0 3248 35 00 0 0 43692 0 83866 0 32519 05 00 0 0 163786 0 93148 0 3248
34. ENTS Requirement Payload or Highest Practicable Level era of Assembly Thernal Vacuum 7 M eraras Temperature Humidity Manned Spaces Temperature Humidity Descent amp Landing Temperature Humidity ER ransportation amp Storage Leakage O 4A oa T A T A T A Applies to hardware carried in unpressurized spaces and to ELV launched hardware Temperature cycling at ambient pressure may be substituted for thermal vacuum temperature cycling if it can be shown by a comprehensive analysis to be acceptable This analysis must show that temperature levels and gradients are as severe in air as in a vacuum Applies to flight hardware located in pressurized area Applies to hardware that must retain a specified performance after retum from orbit and is carried in the unpressurized cargo bay Consideration should be given to environmental control of the enclosure Hardware that passes this test at a lower level of assembly need not be retested at a higher level unless there is reason to suspect its integrity Survival Safehold testing is performed on that equipment which may experience non operating temperature extremes more severe than when operating The equipment tested is not expected to operate properly within specifications until the temperatures have returned to qualification temperatures Test required Analysis required tests may be required to substantiate the analysis Test requi
35. HF Transceiver was made due to the communication requirements of the OBC and the Sphinx NG along with the link capabilities The ISIS UHF Transceiver meets these requirements with a UHF downlink and a VHF uplink option The transceiver comes with a multidirectional antenna giving the CubeSat optimal coverage The cubesatshop com sells this and other versions of the transceivers this particular model is rated at 11336 00 31 39 4 3 Thermal Subsystem 4 3 1 COMSOL Analysis Results When performing simulations COMSOL solved for the temperature of the surrounding chamber and the lamp structure This was undesirable for the model since simulations would take more time to solve for geometries that were not part of the CubeSat COMSOL Support was contacted to see if they could shed some enlightenment to help solve this issue In their response they stated that if the walls of the inner chamber were approximated as far away surfaces then the CubeSat structure can be modeled without the vacuum chamber Since the size of the skeleton is comparable to the actual chamber it had to be included in the analysis Based on this feedback simulations were performed to see if adding the chamber to the model would have a significant effect on the final calculated temperature For the purposes of the modeling simple geometry was used To reduce the simulation time the lamp and vacuum chamber were modeled as surfaces instead of solid shapes However when
36. Min Max Mean Warsaw Access StartTime UTCG Stop Time UTCG 1 09 30 8 18 06 1 2 13 18 2 17 18 4 3 46 51 1 57 16 4 4 23 02 6 32 04 0 5 10 48 7 18 11 4 6 44 20 2 54 53 1 Total Min Max Mean Figure 16 Access Report for the Four selected Ground Stations As the access report in Figure 16 shows the total data in a 24 hour period is 98 Mb To Duration sec 632 4 357 7 463 5 632 8 439 8 538 3 3064 7 357 7 632 8 510 7 Duration sec 515 2 240 2 625 3 541 3 442 7 632 8 2997 7 240 2 632 8 499 6 increase the possible amount of data transmitted more research on adding more ground stations to expand the coverage of the satellite will have to be done Figure 16also shows the most the satellite will pass over any given ground station is six times every fifteen orbits Ground stations in the area of Asia or at least further away from other ground stations will need to be researched since there is significant overlap between current ground stations being used 38 3 2 3 STK Analysis A critical step in design of a telecommunications system is modelling of the links through STK an AGI Analytical Graphics Inc software suite designed to model various satellite systems and missions This project made use of various communications tools within STK Scenario STK allows users to define orbital and mission based parameters to simulate various systems and scenarios Various orbits can be simulated to understan
37. OL generates when trying to apply the mesh The option to use a physics controlled mesh was not possible for most of the CubeSat structure When physics controlled mesh is selected COMSOL automatically applies a mesh to components of the 66 model There were many narrow edges that required sizing techniques Figure 31 shows the final mesh for the CubeSat E 1 10 Figure 31 Mesh for External Components To analyze the temperature range during multiple orbits the range of time was defined to be from 0 to 10200 seconds with a 510 second interval 8 5 minutes This interval was chosen to minimize the calculation time without sacrificing the time resolution of the results Because of its fine mesh choosing a smaller number for the time interval would result in the simulation taking hours to run After running the simulation to see if there was a temperature trend over multiple orbits a line graph was created All edges were selected and then plotted over time 67 Temperature K T T T T T T 280F 270 260 F 250 240 230 f 220 0 2000 4000 6000 8000 10000 12000 14000 16000 18000 20000 Time s Figure 32 Temperature vs Time over 3 orbits Figure 32 shows the temperature ranges over the time interval In this plot the vertical lines represent the range of temperatures between the minimum and the maximum calculated for the entire structure 68 5 Conclusions and R
38. Power Subsystem Hardware 4 2 1 3 Power Subsystem Related Interface Control Document ICD Requirements 9 2 2 Telecommunication Subsystem 12 2 2 1 Telecommunication Subsystem Related ICD Requirements 12 2 2 2 Hardware 13 2 2 3 Ground stations 17 2 3 Thermal Subsystem 20 2 3 1 Thermal Control Related ICD Requirements 20 2 3 2 Analyses Required 20 2 3 3 Active Passive Thermal Control Methods 25 3 Methodology 27 3 1 Power Subsystem Distribution 27 3 1 1 Telemetry Communication 29 fis lt 3 1 2 EPS Switch Configuration 3 2 Telecommunication Subsystem 3 2 Implementation of Telecommunication Subsystem Hardware 3 2 2 Ground Station Network 3 2 3 STK Analysis 3 3 Thermal Subsystem 3 3 Thermal Analysis from External Sources 3 3 2 STK Analysis 3 3 3 Vacuum Modeling in COMSOL 3 3 4 Lab Option in Vacuum Chamber 3 3 5 Thermal Control System 4 Results 4 Power Subsystem 4 1 1 Power Subsystem Hardware Monetary and Power Budget 4 1 2 Wiring Results 4 2 Telecommunication Subsystem 4 2 Mission Modeling and Timelines 4 2 2 Communication Hardware 4 3 Thermal Subsystem 4 3 COMSOL Analysis Results 4 3 2 Thermal Analysis from External Sources 5 Conclusions and Recommendations 5 Power Subsystem 5 2 Telecommunication Subsystem 5 2 Hardware Testing Recommendations 5 2 2 WPI Ground Station Proposal 5 3 Thermal Subsystem 5 3 Recommendations Based on Vacuum Thermal Analysis 5 3 2 Lab Option 5 3 3 Thermal Control System Works Cited 30
39. Project JB3 CBS3 Thermal Telecommunication and Power Systems for a CubeSat A Major Qualifying Project Submitted to the Faculty of WORCESTER POLYTECHNIC INSTITUTE in partial fulfillment of the requirements for the Degree of Bachelor of Science in Aerospace Engineering By Jennifer Hanley Brian Joseph Martha Miller Samantha Monte Joshua Trudeau Racheal Weinrick April 25 2013 Prof John Blandino Project Advisor Abstract The objective of this project was to design the power telecommunication and thermal control subsystems for an earth orbiting CubeSat This mission payload is an X ray detector designed to study solar radiation Requirements on the spacecraft imposed by the National Aeronautics and Space Administration NASA and California Polytechnic State University Cal Poly were reviewed and organized to provide a reference for future design teams The power subsystem defined by previous Worcester Polytechnic Institute WPI student projects was re evaluated and the power budget finalized In addition wiring diagrams were created to show how the power subsystem hardware interfaces with other spacecraft systems The telecommunication subsystem was designed in order to allow communication between the satellite and ground stations A ground station plan was established including a cost budget for hardware and identification of an existing network which could support the mission objectives With this information a telecommu
40. Radio Regulations Article 1 Definitions of Radio Services retrieved 2009 04 23 13 J Wertz and W Larson Communications Architecture in Space Mission Analysis and Design 3rd edition Microcosm Press 1999 pp 533 586 14 NASA General Environmental Verification Standard GEVS 2005 pp 2 6 1 2 4 15 C D Brown Elements of Spacecraft Design Reston VA American Institute of Aeronautics and Astronautics Inc 2002 16 R S D Fred J Dietrich Communications Architecture in Spacecraft Mission Analysis and Design 3rd ed Microcosm 1999 pp 533 586 17 Cubesat GENSO 2013 Online Available http www cubesat org index php collaborate genso Accessed 2012 2013 18 E Dawson N Nassiff and D Velez Attitude Determination And Control System Design For A CubeSat Mission Worcester Polytechnic Institute Worcester 2012 19 Pumpkin TM CubeSat Kit TM September 2009 Online Available http cubesatkit com docs datashee DS CSK MB 710 00484 D pdf Accessed 3 March 2013 20 Clyde Space CubeSat ADCS Module 2013 Online Available http www clyde space com cubesat_shop adcs adcs_board 299_cubesat adcs module Accessed 3 March 2013 21 Comtech AeroAstro AeroAstro Space Micro 28 December 2012 Online Available http www spacemicro com Comtech Areoastro Data Sheets Coarse Sun Sensors pdf Accessed 3 March 2013 22 Analog Devices Inc Analog Devices A
41. S red 2 7 63 PDM 20 SW 16 8 3 BATT 0 5 7 PDM 27 SW 18 8 3 BATT 1 7 PDM 24 SW 21 opt 2 5 7 PDM 26 5V Board Stack Connection 6 Only 28 3 3V Board Stack Connection 6 Only 30 GND Magnetometer CSS1 4 7 62 21 32 GND Payload Gyro 7 61 34 Batt POS Pull Pin Removed 6 EPS DO NOT CONNECT 36 PCM IN Deployment Switch 1 amp 2 6 EPS input to PCM s and PDM switch on state Power Flow to Board Stack 38 Dummy Pull pin off state 6 Load EPS 40 Unused Connection 6 Deployment switch 1 amp 2 off state 42 BCR Out BCR out to Deployment 6 EPS Switches 44 BCR Out BCR Out To Deployment 6 EPS Switches 46 BATTERY Board Stack Connection 6 Only 48 RS422 TXA 7 50 RS422 TXB 7 116 52 OPTIONAL V 7 Table 11 Device Connection Information Device Pin Use Type of Connection References Number Connectio Wire n Number Wire Color Magnetic Red 5 05V CubeSat bus header 2 7 23 Torquer 1 pin 13 Black Ground CubeSat bus header 2 7 23 pin 29 Magnetic Red 5 05V CubeSat bus header 2 7 23 Torquer 2 pin 14 Black Ground CubeSat bus header 2 7 23 pin 29 Magnetic Red 5 05V CubeSat bus header 2 7 23 Torquer 3 pin 15 Black Ground CubeSat bus header 2
42. Torino Torino Italy University of Michigan Ann Arbor MI United States of America University of Texas Austin TX United States of America University of New Mexico Albuquerque NM United States of America Delft University of Technology Delft South Holland Netherlands University of Applied Sciences Heidelberg Germany University of Montpellier Montpellier France University of Vigo Vigo Galicia Spain Table 2 Preliminary 10 Ground Station Locations To input each ground station latitude and longitude of each ground station was found and recorded This provided a basic ground station model in STK but allowed for further restrictions to be added such as altitude of the ground station A ground track and access a report of coverage could then be created using through STK The access report states how long the satellite is in contact with each ground station and how many times it is in contact 34 n ne amp 4 Cubesst STK 9 20 Graphics 1 Earth ce BN ree Edt Yew jnst Anay Satelte tives Window Hed ata SUF BAS d8i w Sg B X 068 B UUW P IP amp MSD 1 no 2012 17 00 00 000 Object Browser 73 w w maa Ouji Educational Use Only uU LT Rin men Bi2Gans Sateltel 88 602 97 165 16 Nov 2012 17 00 00 000 Tene Step 60 00 sec Figure 13 Ground track with ground stations 35 BOs SMH BAR Gd UHESE 0 5 6 6 p 144i 4 WU P IP B GG Q 1 nov 2012 17 00 00 000 Object B
43. W 19 8 3 4 7 3 PDM SW 22 opt 2 5 4 7 5 PDM AI GND 3 board stack connection only 7 6 7 EPS AI GND 3 board stack connection only 7 6 9 EPS AI GND 3 board stack connection only 7 6 11 EPS 13 E g gt 7 6 15 a g 7 6 AI GND 3 board stack connection only 7 6 17 EPS AI Y CURRENT 3 board stack connection only 7 6 19 EPS Solar Array AI Y 3 board stack connection only 7 6 TEMPERATURE 21 EPS Solar Array AI Y PAIR 3 board stack connection only 7 6 VOLTAGE 23 EPS Solar Array 112 AI Y CURRENT 3 board stack connection only 7 6 25 EPS Solar Array AL Y 3 board stack connection only 7 6 TEMPERATURE 27 EPS Solar Array AI X PAIR 3 board stack connection only 7 6 VOLTAGE 29 EPS Solar Array 3l Al 7 6 33 AT 7 6 35 AT 7 6 37 JAI 7 6 39 AT 7 6 41 I2C DATA 3 board stack connection only 7 6 43 12C CLOCK 3 board stack connection only 7 6 AI GND 3 board stack connection only 7 6 45 EPS AI X CURRENT 3 board stack connection only 7 6 47 EPS Solar Array AI X 3 board stack connection only 7 6 TEMPERATURE 49 EPS Solar Array AI GND 3 board stack connecti
44. acking some satellites Make sure to read the README documents Software Space Track The official Keplerian Elements source Requires free locating registration software AMSAT Keplerian Elements for all the amateur satellites in 2 line locating format software locating Celestrak Another distributor of elements software 102 Appendix 7 Complete Hardware Parts List with Images The table below is a hardware and parts list for the major components required to implement the CubeSat mission This includes on board ADC Power and Communication parts and the ground station hardware and software Group Component Company Part Number Image OBC OBC Tyvak Intrepid 20 ADC board ClydeSpace uds pou UD ComTech CubeSat Sun 30 Sensor 3 or 5 AeroAstro Sensor SSBV Space 9 Pine Sun and Ground Fine Sun Sensor Sensor Systems 22 PA Gyro Analog ADXRS450 Devices ADC 46 Magnetic Zarm Technik MTO 5 0 Torquers 3 AG Optimized 24 Magnetometer Honeywell HMC5883L HMC5883L 103 Surrey Satellite 25 GPS Technology SGR 05U US LLC Payload Instrument Sphinx NG XRAY 6 EPS Board ClydeSpace CS 3UEPS2 NB 7 0035 PDM Board ClydeSpace es T mee Power 26 Battery ClydeS pace CS SBAT2 30 Battery FTDI USB id Charge Cord ie RS422 WE Battery 48 Charge Cord Bravekit FTDI UM232R Adapter Front Solar SP L F2U 27 Panels C1ydeSpace
45. actual orbit there should remain a balance between the location and spread of ground stations as one station loses access and another gains access The scenario modeled is likely a worst case in which the area of higher coverage over North America is traversed less than the area of lower coverage over mainland Europe In all cases simulated a future WPI ground station is a good middle ground with around 2000 seconds of downlink access per day and 2300 seconds of uplink coverage In Table 8 the average access time was used to calculate a mean data throughput per day The main link characteristics are summarized and then the packet protocol is specified and used to calculate an actual data rate The Bit Error Rate BER is also accounted for with a worst case value of 10E 5 used UHF Downlink VHF Uplink Mean 10609 7 s day 12076 1 s day Parameter Value Units Value Units Frequency 435 MHz 135 MHz Transmitter Power 1 Watt 1 Watt Trans Bit rate 9600 bps 1200 bps Duty Cycle 1 1 CubeSat Ant Gain 5 dB 5 dB GS Ant Gain 15 5 dB 14 39 dB Pointing Losses 1 dB 1 dB Desired Signal Noise 15 dBm 15 dBm Noise Temperature 542 K 542 K Receiver Sensitivity N A 100 dBm Packet Protocol AX 25 AX 25 Max Packet Size 256 bytes 256 bytes Packets second 4 69 0 586 Overhead 2 bytes 21 bytes Data per packet 235 bytes 235 bytes Adjusted Data Rate 1101 6 byte
46. ader 2 7 30 pin 10 TBD GND CubeSat bus header 2 7 30 pin 29 119 Appendix 10 On Board Computer OBC 64 This appendix shows the interface connections of the OBC As it is not yet known if the OBC can connect directly to the vertical board stack this appendix can be used as a starting point for future teams to wire the OBC to the vertical board stack Intrepid Pico Class CubeSat System Board Interfaces note unused peripherals act as additional GPIOs Daughterboard A Interface SPI with 1 Select 12C 1 GPIO 1 UART 1 Hardware Interrupt Dedicated resistor adjustable 1 2 to 5 5V Regulator 8W peak power Custom Battery Interface Low power 3 3V for logic and sensors Battery Ground and Unregulated Power Unregulated Battery Power DC Da Hard Reboot Line Processor Power On Reset ughterboard B Interface SPI with 1 Select 12C 1 GPIO 1 UART 1 Hardware Interrupt USB 2 0 12Mbps Dedicated resistor adjustable 1 2 to 5 5V Regulator 8W peak power Low power 3 3V for logic and sensors Unregulated Battery Power Line to force hard reboot Battery Ground for Custom Battery Packs Processor Power On Reset Inter Board Connectors SPI with 12 Selects 212C 1 Dedicated for Payload 5 GPIOs 3 UART USB 2 0 12Mbps CMOS Image Sensor Interface 16 pins 5 0V Regulated Power 8W peak power 3 3V Regulated Power 8W peak power Unregulated Battery Power Deployment In
47. and ground operations This is necessary for two reasons 1 To satisfy the mission requirements for retrieving instrument data 2 Tosatisfy the mission requirements for telemetry tracking and command 13 A CubeSat is only as good as the information it provides to the organizations sending it into orbit Information stored on a CubeSat does no good until it is transmitted to ground stations by some means This is achieved through radios operating under Ultra High Frequency UHF Very High Frequency VHF or S band specifications which communicate with ground networks The full system is composed of two main elements the transceiver which is the radio itself and an antenna that tunes the radio s signal Data handling and compression are also a part of the telecommunications system A representation of the system is shown below in Figure 7 Figure 7 Telecommunications subsystem The main driving factors in specifying telecommunications components are data rates reliability of data delivery and power consumption Data rates are driven by instrument data requirements and are a function of signal sampling rate and bit rates The reliability of the system is based on the robustness of the physical components along with correct specification of frequency bands and compression algorithms Data is transferred by the use of packets which are strings of data including control and user information Packets received from the CubeSat sho
48. and research related to existing CubeSat missions revealed that a CubeSat mission must follow specifications given by the following institutions California Polytechnic State University Cal Poly 2 National Aeronautics Space Association NASA 3 and the United States Air Force USAF 4 The necessary requirements were compiled for future design teams reference Within the power subsystem there was an existing power budget to ensure there will always be enough power supplied to all hardware within the CubeSat As changes were made to on board hardware the budget was updated with the respective required power supply data As the hardware for all other subsystems became finalized the requirements for the power supply system became clearer and therefore helped to finalize the power subsystem hardware The wiring for power supply was defined in detail and shared with the CubeSat Structural team Telecommunications on board and between the CubeSat and ground station were investigated for the first time in the CubeSat Design process Ground stations were researched to determine the best hardware and ground station network for a successful CubeSat mission It was determined that joining the Global Educational Network for Satellite Operations GENSO would be an effective way to become part of an established network in a relatively short time Creating a link and access budget using System Tools Kit STK AGI Exton Pa was crucial in choosing the
49. anges of operable temperatures and the minimum temperature at which the entire system can function is 0 C and the maximum is at 50 C Since the temperature in space can range from 100 C to 120 C 29 this limited range of operational temperatures means that the use of heaters may be necessary There will also need to be insulation to prevent overheating while in direct sunlight 50 4 Results 4 1 Power Subsystem 4 1 1 Power Subsystem Hardware Monetary and Power Budget As the details of the project changed the previously discussed power budget was updated As hardware became finalized power consumption values were established Table 6 lists the different hardware items and their respective average power consumption values at the time this report was completed The table was created in order to ensure that margins of power will always cover the entire system at any point Each phase of the orbit will require different amounts of power supply from each piece of hardware The solar power supply however will fluctuate over the course of the CubeSat s orbit as the spacecraft goes into and out of eclipse During sunlight periods the solar panels will re charge the battery for use during the periods of time in eclipse With Table 6 establishing power by mission phase was simple to budget 51 Peak Nominal Quiescent Group Component Company Part Number Power Power Power Es
50. applying boundary conditions surfaces that are not part of a solid cannot be selected for any of the heat transfer choices such as initial values and heat flux Because of this the chamber was given a relatively thin thickness To be consistent the chamber was defined using dimensions approximating the actual chamber used for testing The chamber has a 24 625 inch inner radius and a length of 70 5 inches For simplicity the CubeSat was modeled as a 10 by 10 by 30 centimeter rectangular prism Figure 26 shows the final model for the vacuum chamber simulation 60 Figure 26 Vacuum Chamber Model Once the geometry was formulated the environment in the chamber was created Using the materials browser materials can be selected and applied to the skeleton and chamber The skeleton is made out of 6061 aluminum The vacuum chamber is made from stainless steel and was modeled as SS 316L The outer walls of the chamber were set with an initial temperature of 250 degrees Kelvin Next using the External Radiation Source module heat flux from the lamp was simulated onto the CubeSat structure The point coordinate option was selected The coordinates of the heat source was positioned at 0 20 0 centimeters with a power output of 250 Watts Surface to surface radiation was applied to the geometry The stainless steel vacuum chamber has an emissivity of 0 8 while the aluminum CubeSat has 0 77 In the surface to surface setting the radiosity can be cal
51. at Flight Identical unit MP 3 dB for 2 minutes each c 3 axes 1 25 x MPE Testing shall be performed for content hat is nct cavered by randem vibrzzan testing MPE 3 dP 1 times In bcth amp reczons af 3 axes MP i C Min mum Range 14 3 0 C to 71 3 C Cycles 4 Dwell Time 1 hcur min 2 exreme Temp after thermal stzhilization Transition 5 C minute Vazcum 1x10 Torr Min Temp 70 C Cycles 1 Dwell Time Min 3 heer after thermal stabilization Transition N A Vacuum 1xi Torr Min Temp 70 C Cycles z 1 Dwell Time Min 3 heer after thermal stabilization Transition 5 C minote Vazcum 1x10 Torr PPOD Fight unt includes fight NEA cable and cannecter CubeSat Flight nt MPE fer 1 mince each of 3 axes MPE Testing shall be performed fer cantent that is nat covered hy random vibraben esang MP ESM S C Mir mum Range 9 3 0 C to 66 0 3 Cycles Dwell Time 1 hcur min extreme Temp aer ermal stehilization Transition 5 C minote Vzcuum 1x105 Torr Min Temp 70 C Cycles z 1 Dwell T me Min 3 hreur after thermal sabilizaben Transition N A Vacuum 1x10 Torr Min Temp FUC Cycles Dwell Time Min 2 heer after thermal sabilizaben Transiten lt 5 C minuze Vacuum 1x10 Torr PPOD Fight unit includes fight NEA cable and cannectrc CubeSat Flight unt Dynamic Environments rand
52. ation California Polytechnic State San Luis Obispo California N6CP University Table 3 Final Ground Station Locations The appropriate antenna information was added to each ground station and a second access report and link report was run From the link report shown in Figure 15 the amount of data that will be transmitted during the 24 hour period is given Time s Data Mb Total 10008 3 96 0 Min 240 2 2 3 Max 632 8 6 0 Mean 500 4 4 8 EIRP dBW Revd Frequency GHz Revd Iso Power dBW Flux Density dBW m 2 g T dB K C No dB Hz Bandwidth kHz C N dB Eb No BER dB Min 4 99 0 43549 145 28 131 04 4 43 87 79 288 3320 4797 1830 Max 23 13 0 43551 118 58 104 34 4 52 114 54 288 5995 7472 1 30 Avg 18 23 0 43549 131 73 117 49 4 49 101 35 288 46 76 6153 1E30 Figure 15 Link Budget 37 Calpoly Access StartTime UTCG Stop Time UTCG 1 08 56 7 19 01 9 2 46 11 5 53 38 8 3 58 224 03 46 3 4 31 58 9 42 227 Total Min Max Mean University of Texas Duration sec 605 1 447 3 323 9 623 7 2000 2 323 9 623 7 500 0 Access StartTime UTCG Stop Time UTCG Duration sec 1 31 11 0 40 47 4 576 3 2 07 40 9 15 48 1 487 1 3 23 32 0 27 52 1 260 0 4 56 32 9 06 54 9 621 9 Total 1945 6 Min 260 0 Max 621 9 Mean 486 4 Access StartTime UTCG S Delft top Time UTCG 1 08 46 6 19 191 2 47 38 8 53 36 5 3 48 26 6 56 10 1 4 23 22 4 33 552 5 00 05 2 07 25 1 6 46 06 8 55 05 2 Total
53. cedure is listed in Appendix 1 Based on analysis prior to testing the temperature of the bakeout can decrease to 60 Celsius However an additional hour is added to the bakeout When completing both vacuum tests the profiles must be the same for consistency So either the 70 or 60 degree Celsius profile must be picked for both tests In the vacuum chamber the pressure should remain constant and not exceed 10 Torr from the original pressure If this happens additional thermal baking is required until the vacuum chamber pressure is stabilized In order for the P POD to pass the Thermal Vacuum Cycle test its maximum temperature range must be MPE 5 C 2 3 3 Active Passive Thermal Control Methods The thermal control system of a spacecraft regulates the temperature for the entire unit This system ensures that the temperature of the spacecraft does not exceed the survivability limits of any component at any time It also ensures that the operational temperature for any component is not exceeded while the component is in use If the temperature is not controlled it can lead to component failure or even the mission The thermal environment in space is extremely cold when there is no direct sunlight radiation or another source of light When the spacecraft is exposed to radiation or sunlight the temperatures can climb above the operable or survivable limits This is dependent on orientation of the CubeSat or position in the orbit However
54. culated COMSOL calculates the radiosity of the surface using Equation 2 J 1 6 e0T Equation 2 61 G is the unknown incident radiation T is the initial temperature of the surface o is the Stefan Boltzmann constant 5 67 x 1078 and e is the emissivity of a material In the m sK y Surface to Surface Radiation setting T and are set as initial values COMSOL calculates G values from the geometric surfaces in the model After completing the simulation results from the two studies were compared qualitatively using two 1D plot groups The first group graphed the temperature over the X Y and Z coordinates of the CubeSat Since the CubeSat is fixed at the origin and its dimensions are 10 by 10 by 30 centimeters the coordinates for the X and Z directions were from 5 to 5 centimeters The coordinates in the Y direction is 15 to 15 centimeters In Figure 27 Temperature vs Coordinate X Y Z the dashed lines represent the study where the vacuum chamber is modeled and the solid lines are from the study where just the CubeSat was modeled There is a significant difference in temperature about a 4 degree increase in temperature range The second group shows the radiosity over the coordinates of the CubeSat Because the radiosity increased the temperature will also increase Figure 28 Figure 29 and Figure 30 show the radiosity over the X Y and Z coordinates of the CubeSat respectively 62 Temperature K Surface ra
55. cuum bakeout must be performed on fully integrated flight CubeSats before integration into the P POD 2 1 How to Test 2 1 1 Please read and understand all steps before performing any actions 2 1 2 Clean the external surface of all hardware with lint free wipes and lab grade isopropyl alcohol before inserting the hardware into the thermal vacuum chamber 2 1 3 Place the clean hardware into the thermal vacuum chamber 4 2 1 4 Bring the chamber to a vacuum level of at least 5 x 10 Torr Outgassing will be easily observed at higher vacuum levels 2 1 5 Record the initial pressure level and temperature in the Bakeout Compliance Checklist attached at the end of this document 2 1 6 Starting at room temperature approximately 25 C raise the temperature of the shroud or heating element to 70 C 2 1 7 As temperature is increasing record the pressure level along with the corresponding temperature every 20 minutes 84 2 1 8 Wait until the exterior surface of the hardware has reached 70 C Note There may be an initial increase in pressure this is to be expected 2 1 9 Let the hardware bake at 70 C for one hour 2 1 10 Record the temperature and pressure every 10 minutes during this first bake making note of any unusual pressure readings 2 1 11 If you do not wish to bring your flight hardware to 70 C you may set the upper temperature extreme to 60 C However you must let the hardware bake for two hours 2 1 12 Brin
56. d the direct effects of orbital characteristics on the system Objects such as satellites ground stations targets and celestial bodies can be added to the model and constrained to the mission specifications The satellite capabilities can be constrained through specification of transmitters receivers and antennas to map the various links Ground stations are indicated on the earth map with the analysis using the transmitters receivers and antennas specified by the user for each of the sites Figure 17 below shows the ground tracks of the CubeSat and locations of the four chosen ground stations in Warsaw Delft Cal Poly and University of Texas The access to these stations is highlighted and color coded to the specific stations Figure 17 Ground tracks of the CubeSat plotted using STK 39 Limits and Constraints The satellite itself is constrained by the orbital parameters which are the baseline orbital characteristics for this year s project a 24 hour period an inclination of 98 44 an altitude of 700 km and R A A N of 142 252 Optimally a sun sychronous repeating trace orbit would be used but this is beyond control The satellite is assumed to have a transmitter operating in the UHF band with the following characteristics Frequency 435 5 MHz Power 0 dBW 1 Watt Data Rate 9 6 Kb s Modulation FSK Pointing loss 1 dB The assumed antenna on the CubeSat is designed for the UHF band with the following
57. dicator Solar Power Processor Power On Reset Umbilical A SPI All Device Ethernet RS232 Terminal Access Two USB 2 0 Host USB 2 0 Device Unregulated Battery Charging Line to force Hard Reboot Processor Power On Reset Umbilical B 212C 3 UART CMOS Image Sensor Interface 6 GPIOs ICE JTAG Debug Interface Payload Programming 3 Pins Four Power Good Indicators for Primary Regulated Power 120
58. diosity Wm X CubeSat Only Y CubeSat Only Z CubeSat Only X With Vacuum Chamber Y With Vacuum Chamber 230 Z With Vacuum Chamber 234 232 228 226 224 222 220 218 216 214 212 210 208 14 12 10 8 6 4 2 0 2 4 6 8 10 12 14 X Y Z Coordinates cm Figure 27 Temperature vs Coordinate X Y Z 240 230 220 210 200 190 180 170 160 150 140 130 120 110 100 90 80 x oordinate cm Figure 28 Radiosity vs Coordinate X 250 X CubeSat Only X with Chamber 63 Surface radiosity wim Surface radiosity wim 250 240 230 220 210 200 190 180 170 160 150 140 130 120 110 100 90 60 250 240 230 220 210 200 190 180 170 160 150 140 130 120 y coordinate cm Figure 29 Radiosity vs Coordinate Y 0 z coordinate cm Figure 30 Radiosity vs Coordinate Z Y CubeSat Only Y with Chamber 64 4 3 2 Thermal Analysis from External Sources From the preliminary analysis in Section 3 3 1 COMSOL can easily solve time dependent solar flux equation The next step was to take the solar flux equation and apply it to the model of external components COMSOL has the ability to import 3 D CAD assemblies to create the geometry With the help of the current Mechanical Subsystem MQP team a simplified structure of the CubeSat model i
59. e The 2011 2012 MQP group identified the Clyde Space PMAD modules as the ideal boards for CubeSat power needs 1 The EPS from Clyde Space has flight heritage in a robust Three Unit 3U CubeSat specific package that can handle the planned solar array configuration and has built in power point tracking It has built in overcurrent and battery under voltage protection and was designed from the start for the CubeSat application An image of the board is shown in Figure 3 Figure 3 Clyde Space 3U EPS 6 PDM Power Distribution Module The PDM distributes power to users along specific power busses which are connectors supplying power and data to all the electronics in the CubeSat It includes the ability to switch different users on and off along with protection for every circuit Power is provided by the EPS whether the satellite is running on battery power alone a mix of array and battery power or the arrays alone The PDM also utilizes overcurrent protection on every circuit to protect users The Clyde Space PDM board is shown in Figure 4 It is designed to integrate directly with the EPS and has the same form factor Figure 4 Clyde Space PDM 7 Battery Batteries supplement the energy output from the solar arrays at peak usage while also providing power while in eclipse Lithium ion or lithium polymer cells are commonly used for their characteristically high energy density compared to alternative cells Unfortunately lit
60. e with each other through an integrated CubeSat bus The CubeSat bus is a 104 pin connector interface that consists of two side by side 52 pin headers that each has a part number of ESQ 126 39 G D and is as labeled in Figure 35 7 Starting from the top a possible board stack is layered as follows the UHF Transceiver the Attitude Determination and Control ADC board the battery board the EPS board the PDM board and the OBC board The UHF Transceiver board regulates the transmitted and received signals The ADC board regulates the attitude of the CubeSat The EPS board converts the power provided by the solar arrays into specific usable voltages and provides power to the battery for charging The PDM board distributes the power to the various components of the CubeSat and the OBC board regulates the entire board stack and all the devices attached to it Electrical power is generated by three solar panels and is routed to the Solar Array SA connectors on the EPS board The SA connectors have a part number of DF13 6P 1 25DSA 50 and are as labeled in Figure 37 6 Each of the three SA connectors allows a maximum of two solar panels to be connected to it enabling a total of six solar panels to be connected to the EPS board 6 SA connectors one and two are for a maximum of eight Watt solar arrays and the third SA connector is for a maximum of three Watt solar arrays 6 As the CubeSat Solar Panels being used each generate more than three Watt
61. e EPS board through the SA connectors labeled in Figure 37 of Appendix 8 and the pin numbering of the SA connectors is as labeled in Figure 38 of Appendix 8 The connections of the solar array wires to the connector pins are listed in Table 11 of Appendix 9 The payload wiring consists of a wire that will provide power through the PDM board a ground wire and a wiring harness that will relay information to the OBC directly The connectors for the Sphinx and OBC are unknown at the present time so the design of the wiring harness to connect them should be a goal of future work The power and ground wire 54 connections to the payload instrument are listed in Table 10 of Appendix 9 Although the connectors to the OBC are unknown the connection types can be found in Appendix 10 The devices in the second group consist of the gyroscope the magnetometer the four Coarse Sun Sensors the CubeSat Sun Sensor and the GPS The devices listed should connect directly to the ADC board from Clyde Space According to one of their representatives the ADC board can be built to suit the interface needs of each of the devices The wires and wire connection points for the gyroscope magnetometer coarse sun sensor and GPS are labeled in Figure 39 Figure 40 Figure 41 and Figure 42 of Appendix 8 The location of the CubeSat sun sensor s wire connection points are unknown and will have to be identified in future work However it is known that the CubeSat sun sensor has f
62. e Temp a amp er thermal Temp after thermal stabilization Temp after thermal stzhilizazan stabilization Transition lt 5 C mtnute Tranzzan lt 5 C minute Transition lt 5 C minute Vacuum ixi Torr Vacuum 1x10 Torr Vacuum 1x10 Torr Min Temp 70 C 4 Min Temp 70 C 47 Cycles z 1 Cycles z 1 Dwell Time Min 3 hour after Dwell Time Min 3 heur after PPOD Only thermal stabilization thermal stakilizatan Ref MIL STD Transition N A Transtan N A 1540 B Vacuum 1x104 Tarr Vacuum 1x10 Torr GSFC STD 7000 Thermal Vac b Min Temp 70 C Min Temp 0 C 5 Bake aut Cycles z 1 Cycles z 1 CubeSat Only Dwell Time Mtn 3 hour after Dwell Time Min 3 heur after Re MIL STD thermal stabilization thermal stahtlizatan 1540 B Transition lt 5 C rntnute Tranzton lt 5 C mtnvte GSFC STD 7000 Vacuum ixit Torr Vacuum 1x10 Torr Dynamic Enviranments random MPE envelapes a P95 50 oc mean 5 dB af fight environments Sinusoidal levels envelope loads predictans and fiightenvtrenments Shock MPE envelaps P35 50 for at least 3 samples with 4 5 dB uncertainty factar applied where less than 3 samples are used Thermal MPE Max predicted via simulatian 11 C for uncertainty Sheck testing ts nat required when the following canditians are met 1 The quallficadon random vihratian test spectrum when converted to an equivalent shock response spectrum 2 sigma response for Q 10 exceeds the qualification shock spectrum requ
63. e University San Luis Obispo California 45 hardware use Chipcon CC1000 Software Yaesu FT 847 radio Radio cc1000 c Software Microchip PIC18LF6720 Chip KPC9612 23 Interface MixW TNC RF Microdevices RF2117 Chip Yaesu 847 radio Radio 2m and 70cm Yagi Antenna HF antenna Antenna Tower Rohn JRM23810 non penetrating roof mount Tower Yaesu G 5500 rotor system Rotator M squared 436CP42 antenna is used for 70cm Antenna SSB Electronics preamplifier Amp LMR 400 coax Cable 50 foot run of 5 8 heliax goes thru Cable 1200 baud can be just straight speaker audio after the filters Interface 9600 baud data must be taken before the filters discriminator tapped Interface CAT interface FT 847 Interface DB 9 serial port needs null modem adapter standard PC Computer operating WindowsXP system MixW TNC COM port drivers MacDopplerPRO Dog Park Software Software two Mac computers MacOS X Computer SatPC32 Software Predict by John A Magliacane KD2BD This is a small terminal based prediction program that will work on nix and DOS platforms It has a very clean and clear interface and will run over an SSH connection It will even run on the built in terminal in MacOS X Software Instant Track from AMSAT This is a DOS only program It can run on current machines but has trouble with the screen If you have an old DOS computer drop this on it and practice tr
64. e devices Header 2 lt 0 o gt Header 1 i Transceiver Figure 35 CubeSat Bus s Pin and Header Labeling 108 ABuuy uBios At BITES UNA sje xeij Solar Array Wire Labeling adapted from 6 Figure 36 z a 4 Figure 37 SA Connector Location and Labeling adapted from 6 109 Solar Array Connecting Side O O O O Figure 38 SA Connector Pin Labeling adapted from 6 ANALOG DEVICES e ADXRS450 e oz lt 25 a ee Zu 3 PDD GND MISO GND CS Figure 39 Gyroscope s Mounting Board Wire Connection Labeling adapted from 61 110 HMC5883L 111 Appendix 9 CubeSat Bus Power Board Pin Assignments This appendix details the connections between the 104 pins of the two headers of the CubeSat bus and devices attached to the CubeSat using the labeling conventions described in Appendix 8 The first two tables describe the purpose of each pin of each header of the CubeSat bus and where they connect to The third table describes the connections between the devices and either the ADC board or the CubeSat bus Table 9 CubeSat Bus Header 1 SWzSwitch Al Analog Input Not Applicable Originating Board Device Pin Purpose Maximum Maximum Connection Reference Voltage Current A S V SW 19 8 3 4 7 1 PDM S
65. e sun is qsun and 7 is the unit normal vector of each face of the CubeSat Q is the total solar flux on the CubeSat Some preliminary analysis was needed to ensure Equation 1 worked First it was tested on an arbitrary Two Dimensional 2D surface of dimensions 10cm by 10cm in order to see the effects of solar flux over time on the surfaces Once the results from the 2D analysis concluded that COMSOL can solve the solar flux equation a Three Dimensional 3D cube of dimensions 10cm by 10cm by 10cm was modeled As shown in Figure 19 there is a significant difference in temperature from the initial to final time step 0 to 172800 seconds With the preliminary analysis complete the analysis was implemented onto the final version of the CubeSat Model 43 Figure 19 Left 3D cube at time t 0 sec amp Right 3D cube at time t 172800 sec 3 3 2 STK Analysis The sun s effect on the CubeSat throughout its orbit was analyzed using a program called Systems Tool Kit This program has the capability to simulate any orbit for any period of time and export the data needed to understand all aspects of the orbit The data gathered for this project were the components of the vector between the satellite and the sun in reference to the spacecraft s coordinate system The data was taken for a simulated orbit over a 24 hour period and the data was recorded every 5 minutes The data consisted of the x component the y component and the z component of the
66. ecommendations 5 1 Power Subsystem The next major step in the power subsystem is to establish a power timeline This timeline is established by the orbit hardware use ADC and instrument use As the CubeSat goes in and out of eclipse it will create a timeline for battery recharging and power supply The users power phases will then be established Each phase will inform when and how much power is necessary Once the phases have been established the OBC transceiver instrument and ADC timeline within each phase can be created From these more detailed timelines an on off sequence can be acquired The power supply cycle will then need to be cross referenced with the power user s requirements to create a complete On Off sequence to ensure sufficient power will be received on time 5 2 Telecommunication Subsystem 5 2 1 Hardware Testing Recommendations Testing of the telecommunications system should be conducted when the hardware is eventually obtained The major focus of testing should be on verification of the manufacturer s specifications comparison to link models verification of link integrity and verification of power system compatibility 32 This can be done through the use of link testing in an Radio Frequency RF isolated chamber to limit any external RF interference with the test The transceiver should be connected to the ground station radio and terminal via a cable with known loss and a variable signal attenuator The
67. en violated and flight termination actions are required TMIG Data as a Metric Tracking Source TMIG data is a mandatory tracking source for pro grams using a launch vehicle inertial guidance system when validated by the AFSPC range or certified to the RCC Standard 324 The TMIG data is required at a 10 samples per second update rate from T O through the end of Wing Safety responsibility An RSD system is required as the primary information display system used by the MFCO to evaluate launch vehicle flight Flight Analysis a part of the Launch Safety Section of Wing Safety shall provide RSD requirements instructions and data necessary for display generation to range personnel responsible for RSD systems operations Real Time Prediction Requirements PP and IIP solutions shall be provided for display on the RSD at a rate of 10 samples per second or faster The PP and IIP solutions shall be computed from data supplied by the tracking sources identified in the RSOR and or its operational supplement as applicable PP and IIP computations and displays shall be single failure tolerant Range specific real time calculated IIP accuracy requirements are provided in AFSPC MAN 91 710 and the SW supplements to AFSPCMAN 91 711 Velocity vs time altitude fast slow time relative to nominal plus count time dynamic nominal and other data identified in the RSOR shall also be displayed Wing unique RSD requirements shall be included in the respective wing
68. erlands 43 The ISIS ground station kit has been designed to be compatible with the Global Educational Network for Satellite Operations GENSO hardware Use 38k4 FSK TNC TNC 12 U 19 rack which allows the ground station to fit in almost any location Tower VHF Yagi antenna Antenna UHF Yagi antenna Antenna VHF Low Noise Amplifier Amp UHF Low Noise Amplifier Amp 21dBic gain S band parabolic reflector Antenna Lightning protection system lightning Heavy duty all weather azimuth elevation rotator Rotator Terminal Node Controllers TNC s for 1200 9600bd AX 25 AFSK FSK and BPSK TNC 38k4 FSK Terminal Node Controller TNC and high speed modification of ground station transceiver option available TNC Uninterruptible Power Supply Power Industrial rackmount PC with Computer tracking software Software FASTRAC University of Texas at Austin Austin Texas 44 hardware use Windows HyperTerminal software MacDoppler for Cocoa software MixW TNC ONEStop software two UHF VHF Yagi antenna IC 910H radio ICOM PS 125 radio Astro Dev Helium 100 Hamtronics r100 VHF FM RECIEVER radio MO 96 9600 baud modem kit modem Kantronics KPC 9612 TNC Hamtronics ta451 modification kit modem modifier Hamtronics r451 modification kit modem modifier Hamtronics crystals tantalum caps 101 Cal Poly Earth Station N6CP California Polytechnic Stat
69. es The Serial interfaces are designed to act as bridges or buffers between serial based devices and the rc network The components of the CubeSat that would connect to the PDM through a serial interface are the magnetometer the GPS and the payload 3 1 2 EPS Switch Configuration The EPS Board has connection points that connect to the two separation switches and the pull pin When the pull pin is inserted in the CubeSat or either of the deployment switches is actuated all power to the CubeSat is stopped A possible connection between the CubeSat bus s second header pins and the deployment switches and the pull pin is shown in Figure 12 below The dummy load prevents damage to the BCR when the solar arrays are attached and the battery is not connected This feature is not a requirement and is only intended as a suggested protection while the pull pin is connected CLYDE SPACE 3U EPS Seperation Switches 1 amp 2 DUMMY LOAD I BATT POS CLYDE SPACE 3U BATTERY Figure 12 Possible Configuration of Pull Pin and Separation Switches adapted from 6 30 3 2 Telecommunication Subsystem 3 2 1 Implementation of Telecommunication Subsystem Hardware For the proposed mission a strong emphasis will be put on transferring data collected by the onboard instrument back to ground stations This requires a telecommunication system structured around multiple robust efficient connections with high data transfers To meet these require
70. ete and the solar panels have been deployed to begin providing power For this period the battery will provide up to 2 4 hours of power to supply the entire CubeSat even in the event that all hardware were to be on at the same time 2 The CubeSat hardware must also be completely enclosed in the skeleton With all of these requirements a method of charging the battery must be established A Universal Serial Bus USB cord and an adapter that can be plugged into the PDM board will allow the battery to be charged to a full 30 Watt Hours in order to begin the de tumble sequence and to deploy the main power source consisting of the solar panels The specifications for the hardware chosen and discussed have been reviewed to insure that the CubeSat can work efficiently The final hardware choices are shown above in the power budget in Table 6 A cost estimate for the hardware can be seen in Table 7 Component Part Number each number cost EPS Board CS 3UEPS2 NB 4 500 00 1 4 500 00 6 PDM Board CN SWT 0035 CS 7 750 00 1 7 750 00 7 Battery CS SBAT2 30 3 550 00 1 3 550 00 26 Front Solar Panels SP L F2U 0033 CS 4 000 00 1 4 000 00 27 Side Solar Panels SP L S2U 0031 CS 4 150 00 2 8 300 00 28 Total 28 100 00 Table 7 Power Hardware Costs All power hardware chosen for the CubeSat is manufactured by the same company Clyde Space The choice to use only one company for all part
71. for the CubeSat Once the current values have been verified and the budget has been updated the flight will then be broken down into the different phases of the mission which will each have their own power budgets based on instrument usage 2 1 2 Power Subsystem Hardware Power is a key element of CubeSat design and can be the determining factor in lifetime of the CubeSat The goal of the power subsystem is threefold 1 To generate and or store power for the CubeSat 2 To condition and distribute power for the CubeSat 3 To protect the CubeSat in the case of a fault In the majority of CubeSats that have been launched the first goal is accomplished with the use of solar arrays in combination with a battery giving the satellite power while in view of the sun or from albedo radiation and retaining power in the battery for peak loads and while in eclipse The second goal is accomplished through the use of power management and distribution PMAD modules These modules ensure that the power is supplied to users throughout the CubeSat and that the power is properly conditioned The final goal is accomplished through the use of regulators and safeguards within the PMAD modules that monitor current flows to prevent damage to other components The PMAD modules should also be able to function if the OBC were to malfunction or need to be rebooted in order for the CubeSat to continuously have power once it exits the Poly Picosatellite Orbital Deployer
72. g the chamber and hardware back to room temperature 2 1 13 As temperature is decreasing record the pressure and temperature every 20 minutes 2 1 14 Keep the shroud and hardware at room temperature for one hour 2 1 15 Record the pressure and temperature every 20 minutes 2 1 16 Bring the shroud and hardware back up to 70 C for the final bakeout 2 1 17 As the temperature increases record the pressure and temperature every 20 minutes 2 1 18 Let the hardware bake for one hour at 70 C 2 1 19 Record the pressure and temperature every 10 minutes 2 1 20 If you do not wish to bring your flight hardware to 70 C you may set the upper temperature extreme to 60 C However you must let the hardware bake for two hours This will eliminate most of the outgassing that will occur at this temperature 4 extreme The pressure should remain constant and should not exceed 1 x 10 85 Torr from the original pressure at room temperature If the pressure does increase longer thermal baking is needed until the pressure stabilizes The shroud and hardware are now brought back to room temperature 2 1 21 Thermal bakeout is now complete 7 Bakeout Profile 70C Bakeout Profile 60C 8 Temperature C 8 0 30 60 90 120 150 180 210 240 270 300 330 360 390 420 450 480 510 540 570 Time min 86 General Environmental Verification Standard GEVS 14 VACUUM THERMAL AND HUMIDITY REQUIREM
73. ground station hardware A monetary budget for the on board and ground station hardware was also created to prepare a Worcester Polytechnic Institute WPI Ground Station proposal The thermal subsystem maintains the spacecraft temperature within acceptable limits The group evaluating this subsystem was responsible for conducting thermal analysis for the CubeSat within the space environment and in the vacuum chamber Spacecraft sun vectors were calculated using STK and imported into COMSOL for external component analysis to generate more accurate results Continuing with the Lab Option from last year s MQP team a procedure of the vacuum chamber test was written A model was generated in COMSOL to compare results from the thermal analysis to the vacuum chamber test xi 1 Introduction 1 1 Project Goals and Objectives The goal of the CubeSat project is to provide a sufficient definition of the spacecraft and mission to support a proposal for a CubeSat Mission to fly the SphinX NG X ray solar flux detector The focus of this team is on three subsystems of the CubeSat the power subsystem the telecommunication subsystem and the thermal subsystem The objectives of the project are to confidently show how each subsystem will work in the CubeSat and provide the necessary analysis and hardware recommendations for the actual systems to be created For the spacecraft to be eventually approved for launch the CubeSat must be able to meet all requirements s
74. he WPI CubeSat but potentially to other satellites in the GENSO network 5 3 Thermal Subsystem 5 3 1 Recommendations Based on Vacuum Thermal Analysis Due to time constraints thermal analysis for the Lab Option could not be completed The experiment that was planned would have used an aluminum model of the structure for testing in 71 a vacuum chamber The aluminum frame should be fitted with panels of aluminum sheets and copper clad PCB to simulate the properties of the body of the actual CubeSat and the solar arrays respectively A 250 Watt halogen bulb should be used to simulate the sun in the vacuum chamber To measure the thermal profile of the model five thermocouples should be placed in strategic positions and the data sent to a LabView file through a Data Acquisition DAQ box A final simulation model for the Lab Option analysis has been created using COMSOL One issue with the model was the mesh was too fine This made the simulation take hours and even days to run Future teams should look into making the mesh a bit coarser to allow relatively quicker simulation times The external components thermal model does not incorporate the rear flux on the solar panels In the original model the rear flux was controlled by the analytic function Modeling of internal components is more complex and requires much more time to analyze Future teams should combine the internal and external components into one model This will ensure an acc
75. hium ion or polymer batteries tend to become unstable near their extreme operating temperatures so it is imperative that the batteries are constantly monitored to avoid loss of the batteries or worse These batteries can only handle a finite number of charge and discharge cycles before failure usually in the range of as few as 50 cycles up to thousands of cycles depending on the battery specification To maximize battery life it is important to properly budget power so the battery never exceeds the recommended depth of discharge which is 20 to 30 for most CubeSat batteries While more battery cells could be added for redundancy and to offset power requirements it should be noted that batteries incur a high cost in both weight and volume Despite these limitations batteries are necessary on most CubeSat missions and play an important role in the power system maintaining power in eclipse and supplementing power at times when need exceeds that produced by the solar panels 2 1 3 Power Subsystem Related Interface Control Document ICD Requirements The requirements outlined in the document supplied by Cal Poly San Luis Obispo 2 stipulate that no electronics shall be active during launch This is to prevent any interference either from electrical or RF sources with the launch vehicle and primary payloads The CubeSat must also have a Remove Before Flight RBF pin The RBF pin when installed must cut all power to the CubeSat bus and must be rem
76. ida ie b eie Pasa ntada 92 Appendix 3 Normalized Beta Values over 24 Hours at 5 minute Step esses 95 Appendix 4 Lab Option Hardware List 14 2 ierit trier e tese acne gen Pea ct d eie dd 96 Appendix 5 WP Ground Sta Ona aniio t tede acct nda Sarena eI Ra Coca tuus bd OL Qe arsi 98 Appendix 6 Other Ground Station Hardware seen 100 Appendix 7 Complete Hardware Parts List with Images see 103 Appendix 8 CubeSat Bus Wiring Diagrams eeeeeeeeseeeeeeeeneen rennen 108 Appendix 9 CubeSat Bus Power Board Pin Assignments essere 112 Appendix 10 On Board Computer OBC 64 sees 120 Executive Summary The CubeSat Design Project began in 2010 since then the design process is well underway working toward funding and an eventual launch in the near future The Mechanical Power and Thermal CubeSat Major Qualifying Project MQP 2012 1 worked to create a preliminary hardware selection and created a model in COMSOL Multiphysics Burlington Ma for internal external and Lab Option simulations This year the project got even closer to reality through hardware finalization and breaking ground in the telecommunication subsystem Due to the cost of such a project and other CubeSat missions already in space there many reference documents specifying requirements for launch and mission operations A thorough review of documentation
77. ifferent places In the event that a cheaper used part is unavailable the Upper value will be available Cost Analysis Upper part use cost ICOM CIV radios IC9100 Radio 3 349 95 Yaesu G5500 rotor 749 95 Yaesu G5500 sep kit 24 00 Rotor Controller GS 232B controller 870 ST 1 Rotor Controller 45 K9612 Kantronics TNC 400 MixW TNC Free K8055 Velleman Switch 58 SP7000 SSB ELECTRONICS amp 412 Ground Station Control Computer comp 1300 1 Terabyte External Hard drive comp 85 Linux Laptops for SDR Control comp 1100 Kantronics 3 plus Modem comp 200 Wireless Network Router BEFSR41 comp 11 Uninterruptable Power Supply 1000AVR power 150 12 Volt Power supply 1692 power 345 RF Directional Thruline Watt Meter Mdl 43 power 350 USRP 2 SDR interface 650 USRP 1 SDR interface 700 USRP daughter board WBX interface 450 USRP daughter board TVRX interface 200 2 Coax protectors Polyphaser IS 50NX C2 Lightning lishtiinyaod 63 suppressors 2 Cable protectors Polyphaser IS RCT Lightning suppressors lightning rod 133 11 646 98 Cost Analysis Lower part use cost per ICOM CIV radios IC910h Radio 2 000 Yaesu G5500 rotor 480 Yaesu G5500 sep kit w above Rotor Controller GS 232B controller 630 ST 1 Rotor Controller 39 K9612 Kantronics TNC 400 MixW TNC Free K805
78. ill be shifted east or west depending on the longitude of the ascending node which vary over time due to perturbations of the orbit If the period is slightly larger than an integer fraction of a day the ground track will shift west over time and will shift east if it s slightly shorter Orbital inclination is the angle formed between the plane of an orbit and the equatorial plane of the Earth The orbital inclination i will range from i to i for the geographic latitudes covered by the ground track The larger the inclination then the further north and south the satellite s ground track will pass An inclination of exactly 90 is said to be in a polar orbit it passes over the Earth s north and south poles 33 To create the CubeSat s ground tracks STK was used to analyze the orbit and input ground station characteristics Baseline orbital characteristics for this year s project a 24 hour period an inclination of 98 44 an altitude of 700 km and relative angle of the ascending node R A A N of 142 252 were used to create a satellite in the program Once the data was entered ground stations could be added Initial studies used ten different ground stations from the Global Educational Network for Satellite Operations GENSO network listed below in Table 2 17 Institution Location Country CalPoly SanLuisObispo CA United States of America Warsaw University of Technology Warsaw Poland Politecnico di
79. irement at all fequencies below 2000 Hz 2 The maximum expected shock spectrum above 2000 Hz dees nct exceed g values equal to 0 8 amp mes the frequency tn Hz atall Fequencies abave 2000 Hz corresponding tn the velocity of 50 inches secand Maximum bake aut temperature set to same maximum temperature for thermal cycle test far consistency assuming bake aut would be performed during same vacuum exposure If the Cube Sat cannet achleve these temperature levels the CubeSat shall hold a minimum temperature of 60 C for a minimum cf heurs Levels are defined to be atthe PPOD ta Launch Vehicle mechanical interface Thermal kake out temperatures are not to exceed qualificidon temperatures Figure 10 P POD and CubeSat Environment Tests Thermal 2011 3 Thermal Balance The purpose of the thermal balance requirement is to verify that the thermal control system is adequate during its orbit For the analysis simulations of extreme hot and cold case environments during orbit are required Creating an analytical model of the CubeSat its components and the space environment allows for analysis of the thermal performance of the spacecraft 14 A model can also predict the thermal performance of the CubeSat in a vacuum 22 chamber Although it is not possible to simulate the exact space environment analysis of the testing environment in the vacuum chamber is practical The Mechanical Power and Thermal 2011 2012 MQP team used COMSOL as a t
80. ironment the CubeSat will encounter should be considered for the two extreme cases where the satellite is fully lit by the sun and when the satellite is in an eclipse 2 3 2 Analyses Required The CubeSat needs to survive in the harsh space environment which depending on the position of the satellite can be either extremely hot or cold The thermal control system needs to keep the temperature within the allowable thermal limits of all components This is why there is 20 a series of mandatory analyses and tests before CubeSat can be launched The GSFC STD 7000 reference document provides a table of requirements that must be met for different levels of assembly as well as identifying if analysis or testing is necessary 14 It also outlines the requirements for each test providing special considerations demonstrations and acceptance requirements The table in Appendix 1 from the GSFC STD 7000 document presents the vacuum thermal and humidity requirements According to the table testing 1s required for all levels of assembly to meet the Thermal Vacuum requirements as listed For Thermal Balance testing and analysis is required for payload or highest level of assembly a completely assembled space craft 14 If analysis verification is provided for subsystems and components testing is not required at these levels of assembly The CubeSat will not have any devices that contain fluids so Temperature Humidity and Leakage analysis and te
81. ith batteries must have ability to receive a transmitter shut down command 2 4 3 RF transmitters gt 1mW must wait 30 min minimum after CubeSat deployment switches are activated from PPOD ejection 2 6 2 10 says 45 min for communication sequences to begin 2 4 4 Need to obtain provide licenses for use of certain frequencies 2 4 4 1 proof of frequency coordination from IARU international amateur radio union 3 2 Inputs for frequency licensing 2 4 5 Orbital decay must be lt 25 years after end of mission life 3 4 1 Processing and Ground Operations Thermal Environments Upon delivery to Cal Poly the P POD team will provide a temperature controlled environment for the CubeSats while in the Cal Poly facilities Temperatures are typically controlled at 75 F 10 F Humidity will also be monitored Expected pre launch environments at the Range are defined as follows Temperature 35 F to 100 F 1 7 C to 37 8 C Relative Humidity 0 to 100 3 4 5 Thermal Environment The CubeSat thermal environments are obtained from integrated thermal analysis ITA The maximum expected temperature range is from 44 7 F to 131 2 F 7 0 C to 55 1 C 3 4 6 Thermal Vacuum Bakeout 80 CubeSats shall perform a thermal vacuum bakeout at a high vacuum level minimum 1x10 4Torr 3 4 6 1 Thermal Vacuum Bakeout Profile The CubeSat shall test to one of the two bakeout profiles outlined in Table 3 and Figure 3 with a temperature ramp rate of no greate
82. lar flux will vary with time The Mechanical Power and Thermal CubeSat Team 2011 2012 defined the magnitude of the solar flux during its orbit by using a rectangle function 1 This function is set with a one variable expression with a lower limit and an upper limit The team set these limits to 0 to represent no sun and 1 for full exposure to sunlight The rectangular function was then imputed into an analytic function where it was periodically repeated for the time of two orbits However the group did not take into account the position of the satellite with respect to the sun That is why the beta values are essential for the 42 thermal analysis It will produce more accurate results in order to create a better thermal management plan To import the spacecraft sun vectors an interpolation function was used as defined under Global Definitions in COMSOL An interpolation function can fill in a table or be imported as csv or txt files A defined interpolation function contains values of t and f t For this model the t represents the time in orbit and f t represents the beta values A csv file was imported for each component of beta Beta will change over time since the position of the satellite is changing throughout its orbit With the beta values defined an equation was created for a time dependent solar flux The solar flux is defined as Ny 7 Bx asun 1367 m2 n H B By nz Q qs m p Equation 1 The average flux from th
83. ls is far from constant and as a result the panels must be optimized for specific temperatures Furthermore as can be seen in Figure 2 solar cells operate more efficiently at lower temperatures and produce a higher peak voltage so heat should be dissipated from the panels as quickly as possible Note that there is an optimal point for a given temperature where power is highest This is the peak power point Temp 0 C Data Temp 25 C Data Temp 70 C Data Temp 85 C Data Temp 0 C Model Temp 25 C Model Solar cell output current A Temp 70 C Model Temp 85 C Model gt gt gt f 0 1 0 2 0 3 0 4 0 5 0 6 0 7 Solar cell output voltage V Figure 2 Solar panel output as a function of temperature 5 The 2011 2012 MQP team chose to use a three panel system for the power generation needs the system consists of a front mounted Two Unit 2U panel and two single deployed 2U panels on the sides as shown in Figure 2 1 EPS Electrical Power System The EPS board controls power handling functions and the power subsystem as a whole It is directly connected to the solar arrays PDM Power Distribution Module battery and OBC and as such is in charge of monitoring output from the arrays power to and from the battery and PDM and supplying information to and taking commands from the OBC It is also tasked with maintaining the solar panels at the peak operating power point for a given temperatur
84. m created were implemented in the model 1 The heat flux was the major change When analyzing the CubeSat in a worst case scenario multiple walls of the spacecraft are subjected to the solar flux If the CubeSat was oriented to have its solar panels facing the sun at all times i e best case scenario only the front face of the spacecraft would be subjected to the solar flux Equation 1 and the earth flux were used to formulate the total heat flux on the CubeSat in orbit shown in Equation 3 Qtot Vearth t Asun nxBx t nyBy t nxBz t Equation 3 In Equation 3 qeartn is the constant earth flux and qsun is the constant sun flux defined in Global Parameters Bx t By t and Bz t are the interpolated functions that are dependent on time The normal unit vectors nx ny and nz are for the X Y and Z faces of the CubeSat As for the remaining boundary conditions surface to ambient radiation was applied to each material This setting defines the ambient temperature and surface emissivity According to several online references the emissivity of a typical multilayer printed circuit board was around 0 85 The emissivity of aluminum stayed the same at 0 77 Ambient temperature represents an equivalent blackbody temperature for the surroundings the vacuum of space that the CubeSat is exposed to or about 30 degrees Kelvin Meshing the simplified CubeSat structure was tedious The main problem with imported CAD files is the errors COMS
85. ments hardware designed for CubeSat communication has been selected to maximize transmitter output and datarates This selection will set the baseline for selecting ground stations to maximize access and in turn the total data that can be transferred UHF Primary Downlink and VHF Uplink The ISIS TRXUV UHF 400 450MH2 VHF 130 160MHz Transceiver has been selected for the primary radio for its reliability and full duplex simultaneous transmit and receive capability It is a well rounded platform offering reasonably high data rates of 9600bits s in a CubeSat specific package designed to work seamlessly with the other components It fulfills all the requirements of a CubeSat transceiver accomodating telemetry downlink tracking and command uplink Modulation is handled by Binary Phase Shift Keying BPSK Frequency Shift Keying FSK Audio Frequency Shift Keying AFSK or Manchester FSK as specified by the user These are simple modulation schemes centered around either phase or frequency shifting to carry data in this case a series of binary bits While there are more powerful modulation techniques to expand bandwidth or raise efficiency compared to these simple modulations schemes the simplicity of the proposed system makes implementation straightforward and makes the system more robust and flexible 3l The antenna options for cubesats are limited due to space requirements Generally CubeSats use one of four configurations monopole
86. mission power achievable at the ground stations compared to the CubeSat transmitter Another factor is the increased distance achievable with VHF compared to UHF at a given power level Overall a mean access of 10609 7 seconds for the downlink and a mean access of 12076 1 seconds for the uplink can be expected over a 24 hour period Figure 24and Figure 25 show the access discussed above divided among the various ground stations 56 3000 2500 2000 Calpoly 1500 Delft University of Texas 1000 Warsaw Worcester Polytechnic Institute 500 Access time per 24 Hour Period seconds Nov Dec Jan Feb Mar Apr May Jun Jul Time of Year 2012 13 Figure 24 Expected daily downlink access at each ground station as a function of launch date 3500 a o a D OAZ 9 2500 9 Delft a _ Mm m s 2000 E FA University of Texas 5 Warsaw 9 1000 E Worcester H 2 500 Polytechnic Institute Ee 0 Nov Dec Jan Feb Mar Apr May Jun Jul Time of Year 2012 13 Figure 25 Expected daily uplink access at each ground station as a function of launch date From the plots there is a preference towards use of the European ground stations in order to maximize total access time This is in part due to the choice of orbit when modeling and may 57 need to be reassessed when the CubeSat s actual orbit is known Although the individual contributions will vary depending on the
87. more than 5 Celsius per minute It must bake for 3 hours to allow The qualifying range is 10 degree increase in maximum and minimum expected flight temperature range This range is determined from the thermal analysis 14 23 proper outgassing of components in the CubeSat Outgassing releases any gases or contaminants that were once trapped frozen or absorbed in the materials If the CubeSat does not outgas properly it could affect the performance of its components The bakeout profile depends on the results from the Thermal Balance analysis For instance the Thermal MPE can help choose a profile that is best for the CubeSat It is important that thermal bakeout temperatures do not exceed qualification temperatures From the plot in Figure 11 there are two profiles each with a different bakeout temperature and duration If there is a reason that the CubeSat cannot test at 70 C the temperature may be lowered to 60 C but it must bake for six hours instead of three 70 1 3 hrs 70 C 6hrs 60 C CubeSat Transition Rate lt 5 C per min 0 100 200 300 400 500 600 700 Time min Figure 11 Thermal Vacuum Bakeout Profile 2 2004 Thermal Vacuum Cycle The Thermal Vacuum Cycle test must be performed on fully assembled CubeSats before integrating into the P POD The DNEPR Safety Compliance Requirements document outlines 24 the procedure for the Thermal Vacuum Cycle test The Thermal Vacuum Cycle pro
88. n SolidWorks was created The model contains the aluminum structure from Pumpkin Inc solar panels and hinges from ClydeSpace and 1 millimeter thick aluminum panels to protect the CubeSat Once the CAD file was imported the material boundaries were applied to the components The Pumpkin structure and hinges are made of 7071 aluminum After looking at the ClydeSpace data sheet the solar panels are actually made of multi layered printed circuit board not solid copper as the Mechanical Power and Thermal CubeSat MQP 2011 2012 team assumed 1 To set up the solar flux equation three interpolated functions were inserted one for each beta component As explained in Section 3 3 1 an interpolated function contains values of t and f t Beta values were calculated from STK and then exported into an Excel file The values were extracted and three txt files were created In each file there were two columns The first column contained the time values from 0 to 172 800 seconds with an interval of 30 seconds In the second column was one component of beta corresponding to each time interval The three interpolation functions were renamed Bx By and Bz Next the solar flux and earth flux constants were defined using the Global Parameters section in COMSOL The constant solar flux is 1367 w and constant earth flux is 245 m m 65 Most of the boundary conditions that Mechanical Power and Thermal CubeSat MQP 2011 2012 tea
89. nalog Devices Inc 2013 Online Available http www analog com en mems sensors mems gyroscopes adxrs450 products product html Accessed 3 March 2013 23 Zarm Technik AG Zarm Technik 2010 Online Available http www zarm technik de downloadfiles ZARMTechnikAG_CubeSatTorquers_web2010 pdf Accessed 3 March 2013 24 Honeywell Sparkfun Electronics Online Available http dInmh9ip6v2uc cloudfront net datasheets Sensors Magneto HMC5883L FDS pdf 75 Accessed 3 March 2013 25 Surrey Satellite Technology US LLC SGR 05U Space GPS Reciever 2013 Online Available http www sst us com shop satellite subsystems gps sgr O5u space gps receiver Accessed 3 March 2013 26 Clyde Space User Manual Standalone 30Wh Battery 28 April 2010 Online Available www clyde space com Accessed 20 February 2013 27 Clyde Space 2U Front Solar Panel 2013 Online Available http www clyde space com cubesat shop solar panels 2u solar panels 87 2u front solar panel Accessed 3 March 2013 28 Clyde Space 2U CubeSat Side Solar Panel 2013 Online Available http www clyde space com cubesat shop solar panels 2u solar panels 79 2u cubesat side solar panel Accessed 3 March 2013 29 I t S Initiative CubeSat Structure University of Illinois Online Available http cubesat ece illinois edu Structure html Accessed 3 March 2013 30 SSBV CubeSat Sun Sensor 13 Febr
90. nications link budget was created and expected ground tracks calculated using Systems Tool Kit STK software To better understand the thermal requirements for the mission calculations of spacecraft sun vectors as a function of time while in orbit were performed using STK This data was then used to simulate such effects on the structure using COMSOL The report concludes with recommendations for thermal vacuum testing and future work with respect to these three subsystems Acknowledgements We would like to thank our advisor for his excellent leadership and guidance throughout our project Professor John J Blandino Ph D Associate Professor Aerospace Engineering Program Department of Mechanical Engineering Worcester Polytechnic Institute We would also like to thank the advisors Professors Gatsonis and Demetriou respectively for their assistance to the overall CubeSat Mission Design Project Professor Nikolaos Gatsonis Ph D Director Aerospace Engineering Program Department of Mechanical Engineering Worcester Polytechnic Institute Professor Michael Demetriou Ph D Professor Aerospace Engineering Program Department of Mechanical Engineering Worcester Polytechnic Institute We also appreciate the collaboration with our peers who worked on the other focus groups of the CubeSat Mission Design Project Structural Team Dylan Raymond Billings Ilea Shaneen Graedel Francis Stephen Hoey Peter Kendall Lavallee Justin Michael Torre
91. nsor s Wire Labeling adapted from 21 sssss 111 Figure 44 GPS s Wire Labeling adapted from 63 esee 111 viii List of Tables Table 1 Final Power Budget from 2012 AE CubeSat MQP 1 esee 3 Table 2 Preliminary 10 Ground Station Locations eesseseseseeeeeeeeenere nennen 34 Table 3 Final Ground Station Locations tei ier ere D tug t des adeste d A a UP e ERU Nee ede 37 Table 4 Lab Option Materials Purchased 7 5 erem retento teet nee aa ee enn Ce ao ina e Pe eun 48 Table 5 Operational Temperat r s s is esner Inte aeter ost A PEE sa Ne RU ed Eua ue vdd gie Pu UL E IR URS RON 49 Table 6 Updated Power Budget oie ee VOR PER SER ER UH NEP SAT SERRE IS QURE A TUE Ce duR 32 Table 75 Power Hardware COSU ieu ioi NE eta Sera ENS HR UR qe ON aaa Vade nu NR ETRAS MESURE 53 Table 8 Key parameters in data rate calculation and expected data values 59 Table 9 C beSat Bus Header Vs vsccc cccccccesesecesscococccessvoccossoctcccesesceesecececdeessbecsesechectcsssversesdencet 112 Table 10 CubeSat Bus Headet 2 2 uec Idee EH ue et EE I tei E EET Eaa 114 Table 11 Device Connection IIformallon u2 thee stot ten e en e esr ba seopdat eared gee 117 1X List of Appendices Appendix 1 Requirements per Reference Documents sese 80 Appendix 2 Definition of Key Terms eite iita e
92. om MPE envelopes a P35 50 or mean 5 d of fight enviranments Sinusaldal levels envelope loads prediczans and flight environments Shock MPE envelops 795 50 far atleast 3 samples with 4 5 dE uncertainty factar applied where less than 2 samples are used Thermal MPE Max predicted via simulaton 11 C far uncertainty Sheck testing Is nct required when the follawing cond tans are met 1 The qual ficotan random vibraticn test spectrum when converted ta an equivalent sheck response spectrum 3 2 gma response far Q 10 exceeds the qualificazan sheck spectrum requirementat all frequencies belaw 2000 Hz 2 The maximum expected shack spectrum shove 2000 Hz does not exceed 3 values equal ta 0 8 times the Frequency in Hz at all frequencies abawe 2000 Ez corresponding ta the velocity of 50 Inches second Maximum bake cct temperature set to same maximum temperature for thermal cycle test far cansistency assuming bake aut would be perfarmed during same vacuum expesure Ifthe CubeSat cannot achieve these temperature levels the CubeSat shall held a minimum temperature of CO C far a minimum af 6 heors Levels are defined ta be at the PPOD to Launch Vehicle mechanical Interface Thermal bake rez temperatures are notta exceed quali icabon temperatures 82 PPOD CubeSat Qualification Acceptance Flow Praia 5 on Od a a Figure 1 PPOD and CubeSat Qualification Acceptance Flow Diagram 6 3 7 PPOD door release must be designed to accep
93. on only 7 6 51 EPS SW 19 8 3 4 7 2 PDM SW 23 opt 2 5 0 5 7 4 PDM SW 24 opt 2 5 1 7 6 PDM AI Battery Bus 3 7 6 Current 8 EPS AI GND 3 7 6 10 EPS AI GND 3 7 6 12 EPS AI GND 3 7 6 14 EPS AI GND 3 7 6 16 EPS AI GND 3 7 6 18 EPS AI GND 3 7 6 20 EPS 22 AI 5V BUS 3 7 6 113 CURRENT EPS AI 3 3V BUS 3 7 6 CURRENT 24 EPS AI GND 3 7 6 26 EPS AI GND 3 7 6 28 EPS AT 7 6 30 EPS AI 7 6 32 EPS 34 RS222RX 7 36 RS232TX 7 38 serial RX1 7 40 serial TXI 7 42 serial RX2 7 44 serial TX2 7 46 serial RX3 7 48 Serial TX3 7 50 Serial RX4 7 52 Serial TX4 7 Table 10 CubeSat Bus Header 2 SW Switch Al Analog Input Not Applicable Originating Board Device Pin Purpose Maximum Maximum Connection References Voltage V Current A 1 SW 1 3 33 0 5 Magnetometer 7 62 PDM 3 SW 3 3 33 0 5 7 PDM 5 SW 5 3 33 1 7 PDM 7 SW 7 3 33 4 7 PDM 9 SW 7 3 33 4 7 PDM 11 SW 9 5 05 0 5 7 PDM 114
94. ool to simulate the CubeSat within the space environment 1 In the thermal simulation factors such as solar and earth fluxes conductive and insulated interfaces ambient temperature and emissivity of materials should be taken into account For the testing aspect of the Thermal Balance requirement the duration depends on the payload mission payload operating modes and time to reach stabilization Stabilization occurs when control sensors change less than 0 05 Celsius per hour over a 6 hour period 14 The CubeSat passes the Thermal Balance test when the difference between the predicted and measured temperatures is within the qualifying range The test precedes the thermal vacuum test in order to establish temperature goals for the vacuum test For instance the analysis gives a range of temperatures of the CubeSat throughout its orbit From there the Thermal Maximum Predicted Environment MPE can be generated Thermal MPE is the maximum temperature from simulation plus 11 degrees Celsius added to account for uncertainty 14 Knowing this it can be much easier to match a profile for the Thermal Vacuum Bakeout and Cycle tests These profiles are flexible for components that have more temperature sensitive ranges Thermal Vacuum Bakeout During the Thermal Vacuum Bakeout test the CubeSat is cleaned and placed in a vacuum chamber at an initial pressure of 10 Torr Then temperature is then increased from 25 C to 70 C at a rate of no
95. our analog wires which will provide information to the ADC board aside from the power and ground wires 30 The connection of the devices to the ADC board and the CubeSat bus are as indicated in Table 11 of Appendix 9 4 2 Telecommunication Subsystem 4 2 1 Mission Modeling and Timelines The STK model allows for precise estimations of both ground station access and link budgets as described in the methodology The data from these reports was imported into Excel files and organized by simulated mission date range and ground station The plots in Figures 25 27 were created from this data 55 13000 TAZ 11500 soso E NS NA 10500 S N Z N 10000 Downlink 9500 9000 Access time per 24 Hour Period seconds 8500 8000 Nov Dec Jan Feb Mar Apr May Jun Jul Time of Year 2012 13 Figure 23 Total expected daily access time as a function of launch date Figure 23 shows the access time that can be expected over all 5 selected ground stations for both the uplink and downlink The data in the figure was generated using data from various simulations with different launch dates in STK The propagation of the CubeSat s orbit is the main determining factor in access time achieved while weather and other atmospheric effects contribute to varying path losses It was expected that the uplink access would be higher than the downlink access in simulations about 2046 on average for downlink due mainly to the increased trans
96. oved after the CubeSat is integrated with the P POD If a RBF pin is not present the satellite must launch with its batteries fully discharged The CubeSat must also have at least one deployment switch located on the z face of the CubeSat as shown in Figure 5 The deployment switch must keep the satellite powered off while it is actuated Once the satellite is deployed the deployment switch will no longer be actuated and this will start the deployment timer The deployment timer requirement is a separate requirement from the deployment switch It ensures that no mechanical structures or appendages are deployed until 30 minutes after the satellite is ejected from the P POD It also ensures that no transmissions are generated until 45 minutes after the satellite is ejected from the P POD The deployment switch must also be able to reset the deployment timer if the switch is toggled from the actuated state and then actuated again After the CubeSat is loaded into the P POD it can then be charged through access ports located on the P POD as shown in Figure 6 However the total stored energy in the battery after charging must be no more than 100 Watt hours As the CubeSat has a battery onboard the CubeSat must also be capable of receiving a command to shut down the transmitter Bl cerLoyment switch STANDOFF IB SEPARATION SPRING STANDOFF OPTION A OPTION B SIDE AY pm California Polybechine State University Sa See ater 7 CUBESAT
97. ower budget based on the instruments selected and the information available at the end of the project Table 1 below shows the final power budget proposed by the 2011 2012 teams Component Part Number Peak Powe mE Magnetorquers 3 Zarm Technik AG MTO 2 1 Optimized ComTech AeroAstro Gyro ADXRS450 Magnetometer Honeywell HMC5883L Instrument crc PDM Board ClydeSpace Total Total Available Y 0 300 0 420 0 050 0 030 0 001 8 000 0 1 0 160 0 180 2 000 11 321 12 520 0 300 0 420 0 050 0 030 0 001 8 000 0 1 0 160 0 180 2 000 11 321 12 520 0 000 0 420 0 050 0 030 0 000 1 000 0 1 0 160 0 180 0 000 2 020 12 520 Nominal Quiescent Current Voltage V dU Power W Power W mA Oo 0 09 X 5 MEN TBD 28 95 0 25 TBD TBD TBD TBD TBD Table 1 Final Power Budget from 2012 AE CubeSat MQP 1 TBD TBD TBD This was a preliminary power budget that does not include all necessary information to properly estimate available power Even though the numbers in Table 1 are based on data from the manufacturer these numbers will be tested by future groups once the hardware is acquired These tests will verify the hardware to make sure they do not deviate from manufacturer values For values listed as To Be Determined TBD a manufacturers value will be inserted and verified to make certain there will be enough power
98. performed the electrical functionality of the satellite needs to be re verified 3 NASA requires the CubeSat also have an End of Mission EOM plan 8 Prior to the EOM the satellite needs to be pacified for earth orbit or earth reentry The passivation of the satellite will entail the removal of all forms of stored energy It must be depleted to a point where it would be insufficient to cause a breakup of the satellite The electrical systems specifically the batteries and the charging circuits shall be part of the passivity analysis 2 2 Telecommunication Subsystem 2 2 1 Telecommunication Subsystem Related ICD Requirements In recent years CubeSat launches have become more common With this increase the potential problems that may be encountered have become more clear but also preventable Requirements for sending a CubeSat into space have been established and continuously updated by a variety of authors For the 3U CubeSat under study for this project the required specifications are given by Cal Poly NASA and the USAF A full list of CubeSat applicable mandates can be found in Appendix 1 These documents create a profile for the size and shape of a CubeSat along with restrictions to prevent any perturbations to the Launch Vehicle LV or I2 other satellites already in orbit There is an orbital verification process which reviews the mission flight plans in order to prevent collisions and ejection conflicts Frequency restrictions
99. r than 5 C per minute Duration Minimum Temperature ELaNa CubeSat T Vac Bakeout Profile ao 4 70 C for 3 Hours or 60 C for 6 Hours 3 hrs 70 C d4 amp hrs P ht ga 241 E Temperature C qe lt Li 100 200 300 400 500 400 Foo Time min Each profile is one continuous soak at the temperature and duration specified 6 2 1 CubeSats shall be designed and verified to the environments defined in Table 1 PPOD and CubeSat Test Environments Testing Table and per Figure 1 PPOD and CubeSat Qualification and Acceptance Test Flow Diagram 81 Table 1 PPOD and CubeSat Environments Test Table vibration CubeSat and PPOD Ref Mil Std 1540C Vibration CubeSat and PPOD Ref Mil Std 1540C Shock CubeSzt and PPOD Ref Mil Std 1540C Ref MIL STD 1540 B GSFC STD 7000 Thermal Vacuum Bake out POD Only Ref MIL STD 1542 B GSFC STD 7000 Thermal Vac Dake out CubeSat Only Ref MIL STD 1542 B GSFC STD 7000 MPE 6 di Testing shall be performed for coment that Is not covered by randam vibration testing MPE 6 d3 3 times in both directions of 3 axes FB 1 1 6 Dispo 3 MPE 10 C Minimum Range 14 3 40 C ta 71 Of PC Cycles 9 Dwell Time 1 haur min 6 exceme Temp afer thermal stabilization Transition lt 5 C mincze Vacwum 1x10 Torr PPOD Flight identical unit Inclodes NEN cable and connectar CubeS
100. red if analysis indicates possible condensation Test is not required at this level of assembly if analysis verification is established for non tested elements Note Card level thermal analysis using qualification level boundary conditions is required to insure derated temperature limits for example junction temperature limits are not exceeded 87 STANDARD MATERIALS AND PROCESSES REQUIREMENTS FOR SPACECRAFT National Aeronautics and Space Administration Washington DC 20546 0001 3 4 1 1 4 Commercial Off The Shelf COTS Hardware a A procedure shall be established to ensure that all vendor designed off the shelf and vendor furnished items are covered by the M amp P requirements of this document b The procedure shall include special considerations for off the shelf hardware where detailed M amp P information may not be available or it may be impractical to impose all the detailed requirements specified in this standard c The procedure shall include provisions for ensuring that this hardware is satisfactory from an overall M amp P standpoint NASA Procedural Requirements for Limiting Orbital Debris w Change 1 5 14 09 limiting orbital debris 8 3 1 2 3 3 1 3 3 3 2 5 The NASA Program Project Manager with the NASA Center SMA organization shall track and monitor the noncompliances to this NPR and NSS 1740 14 or NASA STD 8719 14 as applicable per paragraph P 2 4 with the design and operations of the
101. ree Unit ADC Attitude Determination and Control AFSK Audio Frequency Shift Keying AGI Analytical Graphics Inc Bandwidth The maximum change in frequency for a continuous set of frequencies BCR Battery Charge Regulator BER Bit Error Rate The number of altered or incorrect bits caused by interference distortion or noise divided by the total number of bits in an interval of time BPSK Binary Phase Shift Keying Form of phase modulation where the signal phase can be one of 2 possible values C N Carrier to noise radio dependent on bandwidth C No Carrier to noise ratio independent of bandwidth CAD Computer Aided Design Cal Poly California Polytechnic State University CDMA Code Division Multiple Access Multiple Access technique which uses a PN code to spread a narrow band signal over a wider band COTS Commercial Off The Shelf DAQ Data Aquisition DCGS Delft Command Ground Station DNEPR Russian launch vehicle Carried 14 CubeSats but failed to launch Eb No Receiver bit energy to noise ratio EHF Extremely High Frequency EIRP Effective Isotropically Radiated Power The amount of power that a theoretical isotropic antenna would emit to produce the peak power density observed in the direction of maximum antenna gain Measured as the transmitter power plus the antenna gain minus line losses EoC End of Charge 92 EOM End of Mission EPS Hlectrical Power System FCC Federal Communications Commission
102. rom the instrument is usually carried out by the OBC with the option of adding additional processing units and flash memory for dedicated high data operations 2 2 3 Ground stations An integral part of CubeSat communications are ground stations and ground station networks A ground station is an on ground terminal that links to the CubeSat through antennas 17 transmitter receiver and control equipment to transmit and receive messages track or control the satellite 11 Ground stations are located all around the Earth in a series of networks or within the atmosphere such as using other satellites to communicate There are different uses for ground stations telecommunication with satellites communication with space stations or space probes or tracking The earth stations use radio waves in super high frequency or extremely high frequency bands with a frequency band of 3 GHz to 30 GHz for Super High Frequency SHF and 30 GHz to 300 GHz for Extremely High Frequency EHF 12 When the station successfully transmits these radio waves a telecommunications link is established A link is the communications channel that connects two or more communicating devices There are several different types of telecommunication links uplink downlink forward link and reverse link An uplink is the transmission between the earth terminal and the satellite It is the inverse of the downlink which is the link from the CubeSat to the ground station A
103. round sensor information In a case where six solar panels are used this would total 32 analog inputs and if only three solar panels are being used only 25 of the allotted 32 analog inputs would be used The 32 analog inputs are located on the Header 1 H1 header of the CubeSat bus as shown in Table 9 of Appendix 9 The analog inputs begin their numbering with 9 instead of 1 because the PDM board has a different connector for its own separate eight analog inputs The board interfaces with the 32 analog inputs through one multiplexer and the other 8 analog inputs through another multiplexer This is done because the 32 analog inputs all range from 0 to 3 Volts and the other eight analog inputs may not have the same voltage range The multiplexer is used to forward the selected analog input to the OBC Each of the 24 switches can be commanded on or off by the OBC The status and current through the switch can also be retrieved at any time by the OBC To control the switches and access any of the analog inputs an Inter Integrated Circuit PC network is used An PC network consists of a two line system between the boards a clock line and a data line Each board connected to the CubeSat bus also has an C node on it which can communicate with the master IC node on the OBC The I C network can also be used to specify an initial switch configuration on power up of the CubeSat 29 The PDM Board is also capable of connecting to five serial based devic
104. rowser vax fai Warsaw a ET BEA dis Facity Abuquerque Gals ala BS B Jump To Top Start 16 Nov 2012 17 00 00 000 UTCG Stop 17Nov 2012 17 00 00 000 UTCG Educational Use Only jSatellite Satellitel To Facility CalPoly Facility Delft Facility Heidel Facility Montpellier Facility Torino Facility U Satellitel To CalPoly Global Statistics Min Duration 4 Max Duration 3 Mean Duration Total Duration Satellitei To Delft Global Statistics Min Duration Max Duration Mean Duration ITotal Duration ET Satellitei To Heidel Start Time UTCG 16 Nov 2012 17 34 38 180 17 Nov 2012 02 58 09 974 17 Nov 2012 04 34 58 785 17 Nov 2012 06 18 32 373 17 Nov 2012 14 58 27 099 17 Nov 2012 16 37 40 705 17 Nov 2012 06 18 32 373 17 Nov 2012 04 34 58 785 Start Time UTCG 16 Nov 2012 18 47 24 917 16 Nov 2012 20 25 03 142 16 Nov 2012 22 06 37 753 17 Nov 2012 05 00 27 530 17 Nov 2012 06 38 57 299 17 Nov 2012 08 18 28 288 17 Nov 2012 09 59 50 337 17 Nov 2012 09 59 50 337 16 Nov 2012 20 25 03 142 Stop Time UTCG 16 Nov 2012 17 44 01 505 17 Nov 2012 03 08 39 378 17 Nov 2012 04 49 11 571 17 Nov 2012 06 25 59 575 17 Nov 2012 15 11 37 426 17 Nov 2012 16 50 53 849 17 Nov 2012 06 25 59 575 17 Nov 2012 04 49 11 571 Stop Time UTCG 16 Nov 2012 18 59 31 382 16 Nov 2012 20 39 22 804 16 Nov 2012 22 18 17 513 17 Nov 2012 05 09 52 722 17 Nov 2012 06 52 57 528 17 Nov 2012 08 31 5
105. s ADC Team Assaad T Farhat Jighjigh Tersoo Ivase Ye Lu Alan Thomas Snapp We would also like to thank the correspondents who have assisted us in solving software simulation problems Lei Chen Ph D COMSOL Inc ii Authorship Our project team was divided into three subsystems Thermal Telecommunication and Power The team of Jennifer Hanley Martha Miller and Joshua Trudeau supplied sections regarding the thermal subsystem Brian Joseph Samantha Monte and Racheal Weinrick created the telecommunication sections Brian Joseph Joshua Trudeau and Racheal Weinrick were also the authors of the power sections However due to the collaborative nature of the MQP all members were involved in the editing and revision of the project We certify this final report can be considered a group effort with multiple partners collaborating on each section Jennifer Hanley Brian Joseph Samantha Monte Martha Miller Joshua Trudeau Racheal Weinrick iii Table of Contents Abstract i Acknowledgements ii Authorship iii Table of Contents iv Table of Figures vii List of Tables ix List of Appendices X Executive Summary xi 1 Introduction 1 1 1 Project Goals and Objectives 1 1 2 Power Subsystem Objectives 1 1 3 Telecommunication Subsystem Objectives 4 1 4 Thermal Subsystem Objectives 2 2 Background 3 2 Power Subsystem 3 2 1 1 Mechanical Power and Thermal CubeSat MQP 2012 1 Final Budget 3 2 1 2
106. s only two of the three SA connectors on the EPS board are suitable for connecting to the solar arrays The third SA connector will therefore remain unconnected Each SA connector then connects to its own Battery Charge Regulator BCR on the EPS board The BCR s charge the battery and have two modes of operation The 27 Maximum Power Point Tracking MPPT mode is enabled when the voltage of the battery falls below a preset voltage Once it falls below this voltage the BCR s operate at the maximum power of the solar panels to charge the battery The End of Charge EoC mode is enabled when the voltage of the battery reaches the preset value Once it reaches the preset voltage it is held constant and a current from the solar panels is used to finish charging the battery The BCR s are also responsible for supplying power to two Power Conditioning Modules PCM s The PCM s then condition the power into separate 3 3 Volt and 5 Volt power outputs with a deviation of 1 percent 6 The output power of the PCM s as well as a direct connection to the battery are then routed to the PDM board through the CubeSat bus The PDM board has 24 switches which can be used to turn different components of the CubeSat on or off Switches 1 7 have a max voltage of 3 3 Volts switches 8 14 have a max voltage of 5 Volts and switches 15 19 have a max voltage of 8 3 Volts as they connect directly to the battery Switches 20 24 allow for a user specified max voltage
107. s are relatively small in comparison to the other geometries there is no need to model them in COMSOL A 1 D plot group was chosen to view the temperature readings from the thermocouples In this setting edges or points of the skeleton are selected to view the temperature range for each time interval These plots allow for comparison of results from the analysis to the lab option test 3 3 4 Lab Option in Vacuum Chamber To validate the COMSOL thermal model of the CubeSat structure an experiment in a vacuum chamber was prepared All materials were purchased so but there was not enough time to run the experiment Material Manufacturer Part Number Halogen Lamp and Bulb Workforce 778 980 Aluminum Sheet MSC Industrial Supply Co 09426057 PCB Mouser Electronics 590 588 Thermocouples Omega 5SRTC GG K 24 36 Thermocouple extension wires McMaster Carr 3870K32 Connectors for the thermocouples McMaster Carr 3869K48 to the chamber feedthroughs Chamber electrical feedthrough Kurt J Lesker Co EFT0042031 Table 4 Lab Option Materials Purchased Additional information on each item listed in Table 4 can be found in Appendix 4 The LabView file that will be used to collect and analyze the data was developed by the CubeSat MQP 2011 2012 team that first provided suggestions for a vacuum chamber experiment 1 48 The experimental procedure involves three major steps as follows Prior to actual testing
108. s of this subsystem almost guarantees that the component boards will communicate well between each other 39 4 1 2 Wiring Results The CubeSat bus will connect the transceiver ADC battery EPS and PDM boards together with the possible exception of the OBC as indicated in Figure 34 of Appendix 8 The location and type of the OBC interfaces are as yet unspecified and need to be determined As a result ribbon cable will probably be needed to connect the board stack and the OBC The CubeSat bus has two 52 pin headers which are as labeled in Figure 35 of Appendix 8 The purpose and connections of each of the pins of the CubeSat bus are also listed in Table 9 and Table 10 of Appendix 9 The devices that connect to the board stack can then be divided into three groups The first group consists of devices that use power and relay information to the OBC The second group relays information to the ADC board directly and receives power from the PDM board with the exception of the four Coarse Sun Sensors The third group consists of the magnetic torquers which are actively powered but do not relay any information to the board stack The wire connections to the magnetic torquers via the CubeSat bus are listed in Table 11 of Appendix 9 The devices in the first group consist of the solar arrays and the Sphinx instrument Each of the three solar arrays has three connecting wires as labeled in Figure 36 of Appendix 8 The solar array wires are connected to th
109. s s 137 7 bytes s 8813 bps 1102 bps Access time 10609 7 s 12076 1 s 8 Bit error rate BER The number of altered or incorrect bits caused by interference distortion or noise divided by the total number of bits in an interval of time 58 UHF Downlink VHF Uplink Packets Received 49733 0 7076 6 Data Received 11687248 bytes 1662822 bytes 116 87 Mb 16 63 Mb BER worst case 1 00E 05 1 00E 05 Bit Error 117 bytes 17 Usable Data 11687131 bytes 1662806 bytes 116 87 Mb day 16 63 Mb day Table 8 Key parameters in data rate calculation and expected data values As expected the capability of the UHF downlink is an order of magnitude higher than then that of the VHF uplink This is due solely to the maximum transmitter bit rate as the access time is 20 higher for the VHF link In total the ground stations should be able to receive around 116 87 Mb of data and send 16 63 Mb per day The worst case scenario for bit error is approximately 117 bytes of downloaded data a day requiring either resending of packets or use of methods to fix incorrect bits within packets It should be noted that the 20 increase in coverage via VHF is useful for tracking and establishing a connection with the CubeSat and gives more time to upload important commands 4 2 2 Communication Hardware The only hardware required for this subsystem is the UHF Transceiver The choice of the ISIS TRXUV UHF V
110. signal attenuator is a device that acts as a RF power dissipater reducing the signal strength to incur a loss 69 The transceiver should be set up to send a constant stream of dummy data to the radio Once a connection is established and verified the signal attenuator can be used to mimic various losses The data received can be compared to data transmitted at various loss levels to determine the BER Once the BER reaches a maximum value of 10e 5 the loss should be recorded and compared with values from the STK model This loss counts for the sum of all losses in the system and should be used as a metric to reanalyze the model Figure 33 Basic setup showing ISIS transceiver O 2012 isispace nl Left bench top variable signal attenuator O 2012 jfwindustries com Middle and ICOM radio O 2012 icomamerica com Right 10 33 It is important to evaluate the current needed by the transceiver while running these tests to refine and certify power requirements A digital meter or LabView program could easily be used to monitor current draw by the transceiver when sending and receiving data Also the power consumed while in a beacon mode should be measured and included in the quiescent power budget If the power consumption is higher than expected it may be necessary to run the transmitter only when no scientific data is being collected Beacon Mode allows tracking of the CubeSat when not actively transmitting or receiving packets
111. spacecraft and orbital launch vehicle stages beginning at PDR and shall have the tracking reviewed by the Center SMA organization prior to CDR and launch Requirement57313 When significant capabilities affecting the spacecraft s planned ability to passivate maneuver or reenter at end of life change either through graceful degradation malfunction or via command the EOMP shall be updated annotated by the NASA Program Project Manager Requirement 56867 The NASA Program Project Manager shall ensure that all spacecraft and launch vehicles designed to be operated in GEO are designed to be able to maneuver at least 300 km above GEO altitude closest approach to GEO greater than 300 km above GEO altitude Requirement56882 88 34 3 4 1 3 4 2 3 4 3 Conjunction Assessments during Mission Operations for Earth Orbiting Spacecraft The NASA Program Project Manager shall have conjunction assessment analyses performed routinely for all maneuverable Earth orbiting spacecraft with a perigee height of less than 2000 km in altitude or within 200 km of GEO Requirement56891 Conjunction assessment analyses shall be performed using the USSTRATCOM high accuracy catalog as a minimum Requirement56892 The NASA Program Project Manager shall have a collision risk assessment and risk mitigation process in place for all maneuverable Earth orbiting spacecraft that are performing routine conjunction assessment analyses Requirement 56893
112. spacecraft sun unit vector 44 Figure 20 CubeSat Axes 1 As shown in Figure 20 the x y and z correspond to the CubeSat coordinate system with its origin located in the center of the CubeSat The z axis points in the general direction of the sun not directly at it or there would be no x or y components the y axis points through the top of the CubeSat and the x axis points through the side of the CubeSat The data See Appendix 3 was exported into a csv file to be used to set boundary conditions in the COMSOL program see Section 3 3 1 45 Educational Use Only Figure 21 CubeSat Sun Vector Figure 21 shows the simulated orbit and the vector from the CubeSat pointing to the sun which changes its angle over time It can be seen here that the orbit path goes behind the earth on multiple occasions However the recorded data does not indicate when the CubeSat is behind the Earth or when it is in direct sunlight These periods of occultation need to be independently identified by the user because it will affect the fluctuation of the thermal profile 3 3 3 Vacuum Modeling in COMSOL One of the thermal P POD requirements states that a CubeSat must undergo a Thermal Vacuum Bakeout test before it is integrated into the P POD The Mechanical Power and Thermal CubeSat MQP 2011 2012 team began a lab option to perform a vacuum test using the vacuum chamber located in the basement of Higgins Laboratories on the WPI campus
113. sting will not be required There are two vacuum tests that must be performed before launching the CubeSat Thermal Vacuum Cycle and Thermal Vacuum Bakeout According to the table included in the DNEPR Safety Compliance Requirements reference document the P POD must go through both tests whereas the CubeSat must be tested in the Thermal Vacuum Bakeout 14 For each test the CubeSat and P POD must undergo a protoflight and acceptance test A list of requirements per test is in Figure 10 The components are tested at safe levels to ensure they work DNEPR is a Russian launch vehicle It carried 14 CubeSats to be launched into space but failed to launch 66 Protoflight test levels are the same as qualification levels They are performed on prototype components at extreme levels nearly 1 4 times over the operating loads of the components 65 Acceptance test levels are performed when the CubeSat passes the qualification and protoflight tests 65 21 Table 1 PPOD and CubeSat Environments Test Table Tests Qualification by Test Protoflizht Test Acceptance Test Thermal MPE 10 C MPE 10 C MPE 5 C Vacuum Cycle Minimum Range 14 3 40 Cto Minimum Range 14 3 40 C to Minimum Range 9 3 40 C ta 66 PPOD Only 71 O 4 3 C 471 0 3 C 3 C Re MIL STD Cycles 8 Cycles z 4 Cycks 2 1540 B Dwell Time 1 hour min Dwell Time 1 hour min extreme Dwell Time 1 baur min exreme GSFC STD 7020 extrem
114. supplement to AFSPCMAN 91 711 Video Display Requirements Video monitors with channel switching capability shall be provided to the MFCO and Senior MFCO positions Specific video coverage requirements peculiar to a mission shall be identified in the RSOR for the vehicle or the Ops Sup for the mission Telemetry Display Requirements 90 6 7 6 1 Specific telemetry display requirements shall be documented in the RSOR or Ops 6 10 7 4 1 Sup for a mission Telemetry Systems Real time information on in flight vehicle performance and behavior typically includes but is not limited to engine chamber pressures roll rate attitude launch vehicle velocity vs time automatic gain control values of the command destruct receivers and occurrence of discrete event Additional telemetry system requirements are provided in the SW supplements to AFSPCMAN 91 711 and the program specific RSOR Command Capability AFSPC ranges shall ensure that range managed instrumentation pro vides uninterrupted command capability for all systems that use an FTS or thrust termination system For systems that use command receiver decoders capture from FTS turn on through flight control end of mission is required The Flight Control communication circuit requirements shall be specified in the applicable RSOR Ops Sup or operations directive OD 91 Appendix 2 Definition of Key Terms 2D Two Dimensional 2U Two Unit 3D Three Dimensional 3U Th
115. t redundantly initiated signals 6 4 4 LV shall accommodate PPOD door position indicator in the flight telemetry stream 6 4 5 LV avionics shall provide redundant separation signals to the PPOD door actuation device 6 4 9 LY shall command deployment of the PPOD s CubeSats 64 10 LV trajectory design shall not result in LV contact with deployed CubeSats 6 4 11 LV shall not deploy the CubeSats in a trajectory that will contact the Primary Mission or LV 6 3 15 PPOD must have a fixed base frequency gt 120 Hz 7 Development of Mission Ground Operations Plan a Identification of pre launch activities functional tests handling etc and requirements basic facilities time needed etc 83 b Description of spacecraft operations plan post LV separation particularly transmitter initialization and first contact DNEPR Safety Compliance Requirements 34 2 Thermal Vacuum Compliance Thermal vacuum bakeouts are critical in assemblies of space flight hardware to ensure the lowest levels of outgassing Thermal bakeout of smaller subassemblies and components help reduce overall bake time and decrease the final levels of outgassing A minimum vacuum level of 5 x 10 Torr must be attained to observe the outgassing of components Note NASA certified materials should always be used in space flight hardware especially epoxies and glue If in doubt about the materials you are using please contact the CubeSat Coordinator Thermal va
116. tated by the involved parties These include standards for testing set forth by National Aeronautics Space Association NASA 3 California Polytechnic State University Cal Poly 2 and Department of Defense standards 4 This team was responsible for understanding and creating procedures for these requirements so that the CubeSat Mission could be approved 1 2 Power Subsystem Objectives Power subsystems are used to produce condition store and distribute power to various devices throughout the CubeSat There were two primary objectives for the power subsystem The first objective was to build upon last year s work and have an updated power budget to ensure that the payload and all CubeSat subsystems have the power they need at all times 1 The second objective was to detail the wiring of all the power subsystem circuit boards and the CubeSat devices i e power consumers to ensure that everything was connected and could receive power as well as relay information to the controlling CubeSat On Board Computer OBC 1 3 Telecommunication Subsystem Objectives The CubeSat requires communication between earth and itself Telecommunication subsystems are used to relay commands and data to and from the CubeSat and ground stations on Earth To make sure this was possible for the CubeSat this team was responsible for three primary objectives The first was to determine using Systems Tool Kit STK the extent of ground coverage the C
117. the northern hemisphere A disadvantage of this geometry is that there is a continuous change of antenna positioning and switching links between satellites 13 I9 Geostationary orbit with crosslink is an architecture used when a geostationary satellite is beyond the line of sight of a ground station and a secondary satellite relays data to the ground station A relay satellite is better than using two adjacent ground stations because the adjacent ground station is on foreign territory Since it is on foreign territory there is more cost and the transmission is less secure and less survivable The disadvantage of the relay satellite is it increases the architecture s complexity risk and cost 13 2 3 Thermal Subsystem 2 3 1 Thermal Control Related ICD Requirements A thermal vacuum bakeout test must be performed to ensure acceptable levels of component outgassing 2 Outgassing is the release of volatiles gas or vapor that has been dissolved trapped or absorbed in a material The thermal vacuum bakeout test entails the CubeSat being placed in a high vacuum level of a minimum of 10 Torr and exposed to either 70 C for 3 hours or 60 C for 6 hours The satellite must also be able to survive the temperatures of its environment both during the launch and in space The launch condition temperatures range from 1 7 C to 37 8 C and the thermal environment temperatures during its powered flight range from 7 0 C to 55 1 C 2 The thermal env
118. tructure 10 feet Tower 2 08m parabolic dish Figure 4 with a right hand circularly polarized feed is used to receive the S Band Andersen Manufacturing Antenna Two 435MHz 42 element circularly polarized with polarization switch Yagi Uda UHF signal Antenna single Yagi Uda antenna VHF Antenna S Band signal PSM 4900 by Datum Systems Radio Half duplex Yaesu FT 897 amateur transceiver UHF link Radio VHF radio Yaesu FT 847 Radio Rotator AlfaSpid Radio full Azimuth Elevation design Rotator Custom designed microcontroller uses a simple 8 bit microcontroller uC the A Tmega32 Controller A pair of low cost H Bridge motor Motor 200MHz ARM Single Board Computer SBC by Technological Systems Computer operating Linux system SCC PEB20525 Serial Communications Controller Controller Rotator Interface RS 422 interface Interface S Band Modem Interface Interface Since the TNC control board S Band modem DATUM Systems PS2100 Modem UHF Modem CMX589 synchronous GMSK Modem RS 232 serial ports interface with the UHF radio Interface Plan13 algorithm fed by NORAD TLE s A set of scripts have been developed that interface with the SpaceTrack website to automatically retrieve new NORAD TLE s Software Operator Interface Software created Nanosatellite Interface Control Environment NICE NICE was Software The Terminal Interface Program TIP created Software 100 Delft Command Ground Station DCGS Delft University of Technology Delft Neth
119. uary 2013 Online Available http www ssbv com ProductDatasheets page39 page29 index html Accessed 20 February 2013 31 CubeSatShop com Small Ground Station ISIS 2013 Online Available http www cubesatshop com index php page shop product details amp flypage flypage tpl amp product_id 24 amp category_id 3 amp option com_virtuemart amp Itemid 72 amp vmechk 1 amp Itemid 72 Accessed 2012 2013 32 T Bleier and et al QuakeSat Lessons Learned Notes from the Development of a triple CubeSat Quakefinder LLC 2004 33 JWF JFW Benchtop Rotary Attenuators JFW 2013 Online Available http www jfwindustries com catalog Benchtop_Rotary_Attenuators 61 1 html Accessed 3 March 2013 34 Cal Poly DNEPR Safety Compliance Requirements August 5 2004 pp 2 3 10 35 H DEPOT Workforce250 Watt Halogen Portable Work Light Homer TLC Inc 2011 Online Available http www homedepot com p t 20207 1327 storeId 1005 1 amp langId 1 amp cataloglId 10053 amp productId 20207 1327 amp R 202071327 UTPLwjD_mSp Accessed 2013 76 36 M I S Co Aluminum Sheets MSC Industrial Direct Co Inc 2012 Online Available http www 1 mscdirect com cgi NNSRIT2 7PMAKA 09426057 amp PMPXNO 4463866 amp cm _re ItemDetail _ ResultListing _ SearchResults Accessed 2013 37 M Electonics Copper Clad Boards SINGLE SIDED 6X9 COPPER CLAD 1 32in TTI and Berkshire Hathaway 2013 Online Available
120. ubeSat would have access to through the use of various ground stations The second objective was to define a data link budget based upon the extent of ground coverage as well as the CubeSat telecommunication hardware The third objective was to identify and recommend hardware to be used in the CubeSat Mission 1 4 Thermal Subsystem Objectives The thermal control subsystem ensures the CubeSat does not exceed the maximum and minimum operational temperatures of its components With that in mind this year s team was responsible for two primary objectives The first objective was to refine last year s thermal analysis by using STK to determine the angles at which the sun strikes the CubeSat and then use those values in COMSOL to produce a more detailed thermal modeling of the CubeSat as it orbits the Earth 1 Using this data an analysis can be performed to determine a recommended thermal control method The second objective was to test a model of the CubeSat in the Worcester Polytechnic Institute WPI vacuum chamber for comparison to the COMSOL analysis in order to verify the findings 2 Background 2 1 Power Subsystem 2 1 1 Mechanical Power and Thermal CubeSat MQP 2012 1 Final Budget Power is one of the most important considerations when planning a CubeSat mission If the power budget was incorrect or the power system shorted out the entire mission would be compromised The 2011 2012 Major Qualifying Project MQP team 1 created a p
121. uld be able to be salvaged if incomplete otherwise the packet would need to be resent Frequency ranges for VHF UHF and S bands are as follows VHF 30 300MHz UHF 300 3000MHz S Band 2 4GHz 14 costing precious downlink time Finally power consumption plays a key role as the radio is a comparatively large power draw at up to 1 5W An important aspect of the telecommunications system is the ability of ground stations to monitor telemetry track and command the CubeSat Various telemetry measurements taken by the OBC verify that the CubeSat is functioning properly and if not can give clues as to why a part is malfunctioning Tracking provides verification of the CubeSat s position and trajectory by analyzing the time delay in sent and received packets Finally command functions allow the ground operations to access different modes of the CubeSat or address immediate problems Antenna CubeSat antennas are optimized for small size and low mass while integrating with available UHF VHF or S band transceivers Two common antenna designs for CubeSats are patch antennas which affix to a side of the CubeSat and are little more than a thin flat PCB with an imbedded antenna and deployable tape spring antennas which are affixed to a standard base and spring outwards from recesses in the CubeSat An example of an ISIS deployable antenna system being considered for the current proposal can be seen below 15 Figure 8 ISIS deplo
122. urate thermal control management plan 5 3 2 Lab Option Future teams working on the thermal analysis of the CubeSat model can run the experiment in the vacuum chamber The materials have all been procured and the experiment can now be set up They should also consider designing and running more complex experiments Such experiments could include turning the lamp on and off to simulate the change between eclipse and direct sunlight Experiments could also vary the placement of the thermocouples on the model to see where high or low temperature spots are 72 The results of the vacuum chamber testing are expected to indicate that there is a profile across the skeleton of the CubeSat such that it is warmer on the face pointing at the lamp heat source and is cooler on the back side Another factor that should be considered is the influence of the material of the CubeSat on the temperature The expected observation is that the aluminum surfaces will distribute the temperature to the rest of the CubeSat model more readily than the copper clad PCB The model used in these experiments should be composed of materials with similar properties to those of the external components of the CubeSat The material selection thus far has been working from a very basic assumption that was made for simplicity of data collection and to reduce the number of thermocouples that need to be placed Future teams could consider adding additional internal struct
123. ure to the model to simulate internal components as well 5 3 3 Thermal Control System Once the thermal profile of the CubeSat over its lifetime is finalized and the final structure make up is known the control system options should be determined The factors taken into consideration for the thermal control selection should be primarily concerned with mass and volume limitations of the CubeSat the P POD requirements and the maximum and minimum operational and survivable temperature limits Due to the power restrictions of the CubeSat passive thermal control methods are the only realistic option Future groups will need to look into thermal coatings and ensure that they comply with the P POD requirements Research on passive radiators should also be conducted if there is available space to place such control mechanisms 73 Works Cited 1 J Bauer M Carter K Kelley E Mello S Neu A Orphanos T Shaffer and A Withrow Mechanical Power and Thermal Subsystem Design for a CubeSat Mission Worcester Polytechnic Institute Worcester MA Project JB3 CBS2 2012 2 Cal Poly SLO Launch Services Program Program Level Poly Picosatellite Orbital Delpoyer PPOD and CubeSat Requirements Document National Aeronautics and Space Administration John F Kennedy Space Center Florida LSP REQ 317 01 Rev A 2011 3 National Aeronautics and Space Adminstration Standard Materials and Processes Requirements for Spacecraft Launch Ser
124. vices Program Washington DC NASA STD 6016 2013 4 L C P M Solie Launch Safety Requirements For Air Force Space Comand Organizations Air Force Space Command AFSPCMAN91 711 2007 5 The Mathworks Inc Mathworks 2013 Online Available http www mathworks com products demos shipping elec elec solar opt m 02 png Accessed 3 March 2013 6 Clyde Space User Manual CubeSat 3U Electronic Power System 27 June 2011 Online Available www clyde space com Accessed 3 March 2013 7 Clyde Space CubeSat Power Distribution User Manual 14 March 2011 Online Available www clydespace com Accessed 03 March 2013 8 National Aeronautics and Space Administration NASA Procedural Requirements for Limiting Orbital Debris w Change 1 5 14 09 NASA 2009 9 CubeSatShop com CubeSat Sun Sensor ISIS 2013 Online Available http www cubesatshop com index php page shop product details amp flypage flypage tpl amp product id2104 amp category id 7 amp optionzcom virtuemart amp lItemid 69 Accessed 3 March 2013 10 CubeSatShop com Deployable Antenna System for CubeSats ISIS 2013 Online Available http www cubesatshop com index php page shop product_details amp category_id 6 amp flypa ge flypage tpl amp product_id 66 amp option com_virtuemart amp Itemid 70 Accessed 3 March 2013 74 11 F S 1037C Earth Station General Services Administration 1996 12 I T Union ITU
125. vs Coordinate X Y Z ise RTERS TIRE SANA SUERR S ES AR EESA UE CUAER 63 Figure 30 Radiosity vs Coordinate AC cocos onc tote nacre ek econ a Ato emu ue 63 Figure 31 Radiosity vs Coordinate Y i i reite ceste tate peracta tesa rus diced edu 64 Figure 32 Radiosity vs Coordinate Zeit itae eie toe oh D bes utes Pas orci roten cud ed aues 64 Figure 33 Mesh for Extetnali omponenis sco eene lato ie Natrii vetus 67 Figure 34 Temperature vs Time over 3 orbits 2nd or end teintes fe Peto epic que 68 Figure 35 Basic setup showing ISIS transceiver O 2012 isispace nl Left bench top variable signal attenuator 2012 jfwindustries com Middle and ICOM radio 2012 icomamerica com Rishi PLO 033 Rocco iesu OD WERE M REID estt P stu etat foede vs ub ases Sous rcs de US 70 Figure 36 Vertical Board Stack Side Images of boards adapted from 7 6 and 26 108 Figure 37 CubeSat Bus s Pin and Header Labeling eene 108 Figure 38 Solar Array Wire Labeling adapted from 6 eene 109 vii Figure 39 SA Connector Location and Labeling adapted from 6 sssss 109 Figure 40 SA Connector Pin Labeling adapted from 6 sse 110 Figure 41 Gyroscope s Mounting Board Wire Connection Labeling adapted from 61 110 Figure 42 Magnetometer s Mounting Board Wire Connection Labeling adapted from 62 111 Figure 43 Coarse Sun Se
126. yable antenna 9 The common antenna types seeing use on CubeSats in the aforementioned configurations are monopoles dipoles and turnstiles A monopole takes the form of a single radiating wire with a low gain and nearly spherical radiation profile This makes it favorable for systems with low complexity and little to no pointing ability The dipole and turnstile antenna types are comprised of 2 and 4 monopoles respectively which boosts antenna gain at the expense of more directional radiation profiles Transceiver A transceiver is a single board housing both a transmitter and receiver circuit The transceiver converts data from the OBC into a form that can be sent to ground stations via a carrier signal The most common Transceivers for CubeSats operate in both the UHF and VHF bands This allows the use of the VHF band for downlink and UHF band for uplink or vice Versa 16 Figure 9 ISIS UHF VHF transceiver 10 Data Handling and Storage The CubeSat does not have the luxury of being connected to a ground station during all times the instrument is taking data For this reason it is necessary to include some variety of data handling and storage to process information from the instrument A number of compression algorithms may be applied to the data on board so as to reduce the storage requirement and increase the amount of data transferred to ground stations during the specified fly over time The compression and storage of data f

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