Private Pilot Lectures
Private Pilot Lectures AT 14400
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Date Created: 09/19/15
SATZ tudent Animal Iracking ell e Preliminary Design Review Rough Draft Nat Sangngampal Torn Fosness Kamlesh Nankani Brett Northcutt German Porras Alonso de Celada Index 908994P N Introduction Concept of Operations Orbit Selection Launch Vehicle Integration Spacecraft as a System CommunicationsPayload ADCS CDampH Power Thermal Control Structures and Mechanisms Requirements vs Capabilities Cost amp Future Work 1 INTRODUCTI 1 1M1ssmn StalEmEnl 51171 was mama 5 1 5111111 1513111 andbwh ans 51121111 fax 111 mm armm am 51114 ms 11m 151M 5 m1th ms 5112mm 1 cammzxcul ammal 1mm 11m cumu yuses myhxge expmve 511211125 1 2 Mssmn Objecnves Th2 mun 11mm 11511111511 e wun ymck GP 1111mmquot eqmppd ammals Pasgblz cuslamzxs 111111 scxzm c quotswam liveska awnzxs and wild 211111511515 Ths 11mm 15 m be accumphshzd Wm bung 11w casL simple 51111111 1151 Th2 secundaryabjectm 151a cream a P R 1m fax 111 ME dzpan mzm 11 1mm Unlvexsny 1 3 Satelhte Desmpuun 1 1111mm dummde threwlaw W11 ax swam Nora Dmnaonsam417cm xandydunno and cmmzdvntwn 132 Tatalspcecm mass Th2 spacecm lusa mmmss D172 68 kg 1 z 3 Tan paw Maxxmnm puwens 2 3w memum paw 1511 w Average paw 517W 2 Concept of Operau39ons Fxrst the beam of the animal to be tracked must be determined and then chrs equrpprhg each ahrrhal thh a GPS receiverthat determines its posrhor atngen hme intervals The secondtaskxs aehreved by a small trans mar connectedto a GPS reeerver eh encodes the posrhon reports and once r determines when the warhead from a stored epherh s rurahsrhns the rhforrhahor The n w mm L ephemens Theh r stores the ahrrha posrhor rhforrhahor uhm xtpasses over the Purdue VT F n deT w HM hm m mm The following gure illustrates ths concept Updated Ephemens Acknow edgement Amma SIored Dal Data 5 Te emetry Fxgme 21 SAT operations concept m Mr W eorhrhuhreahorrs strategy can be foundm Seehorrs 3 and s The ehdresuh rs thatthe spacecra rs able to track approximately 12650 ahrrhas per day mh the lxmxtauon of spahhrhg the whole surface othe Earth see Seehor s for detmls 3 Orbit Selection According to the mission objective and requirements the spacecraft must be placed in an orbit that provides worldwide coverage and it is expected that it will be launched as a piggyback payload to an orbit de ned for a primary payload The first requirement limits the spacecraft to near polar orbits and since most primary payloads going to this kind of orbits are Earth observation payloads they tend to go to sunsynchronous repeating ground track orbits that provide an almost constant local time at a given latitude for each ground pass this was the kind of orbit selected Since the orbit is de ned by the primary payload the orbit selected for the purpose this design effort must have as many as possible worst case characteristics from the point of view of the different spacecraft subsystems as detailed in Table 31 Parameter Altitude The higher the altitude the longer the duration of eclipse and sunlight phases leading to higher requirement on batteries and higher temperature oscillations Small increase in radiation damage Lower altitudes lead to shorter view times from ground stations Inclination 139 Main factor affecting radiation damage see Section 8 In general the higher the inclination the more damaging the environment Fixed by sunsynchronous requirement once 139 and e are specified Local time of Descending Node Large impact on eclipse durations Must be around noonmidnight for worst case eclipse conditions Eccentricity e Small impact unless the value is large Typically small for payloads going to sunsynchronous orbits The selected orbit characteristics are described in Table 32 This orbit is very similar to that of Envisat a ESA environment monitoring satellite recently launched by the selected launch vehicle Ariane 5 see Section 4 of this reportThis orbit meets all but one of the characteristics listed in Table 31 and is easily achieved by the launch vehicle The only worst case condxtxon not met 15 that ofhavmg shortvxew umes om the munw nm h vxew he table below starttlmes lxsted Only one pass per day thh a mer 1 day and the commumcauons oppormhmes thh the ground station t L dumuon geater than 1200 seconds 20 Da 1 Da 2 I a 3 Start Durau0hsee Sm Dummhsee Sm Dumuonsee m 53 53 A m sue m 25 he he ADA m 5 5 5 5 717 as m 5 32 5A neA hen m 3 he 55 1sz a7 5 3 55 1215 722 DA A n he 157A 557 he 25 A 2 1557 MA 5 5 5 5 15A 172 he 53 m A 5 5 m 25 A he uzs m7 5 5 55 5 1A3 m A As 5 52 mm m 5 2A B a 1er A57 5 55 5 5 m Am 7 A A2 55 1 ms 3 23 5 us 1572 m 5 5 5 255 52 5 AA 35 A A 577 2 25 DA en 1251 eAe 5 5 5 55 1A5 21 As Ae 35 m3 321 22 25 en 75 as 575 5 m 5 5 Au mu 2 AA he 75 5 5 55 5 5 mA 22 us ze 25 ms 9 A r 9 Fxgwe 3 ypxcd one day ground trackarthe Selected arbzr m quotn ofthe penod whAeh As closeto the Lyplcalmaxlmum for LEO orbxts ofabout40 n T F0 r whAeh results Ah harder power and thermal requirements h we M 3052 ar echse N 37 quotn oforbxtpenod 4 Launch Vehicle Integration The launeh vehtele seleetedts the Anane 5 The reason forthls 15 that of all the major t tertn terms of mass and volume only slx rernatn tn tnventory before rt ls phased out tn 2003 sothe Anane 5 requtrernents wlll be used At a later date rt seems that Boelng tntends to tnstate a slmllar program for the new De ta IV onee rt beeornes n r h w L swltchlll u m one to another shouldbe relaavely stragntforward by changlng adapters w volurne adapterrneehantsrn and dynamte and staae response ofthe spaeeeratt The statte Seetton 11 Table 41descnbes the mass volume and attaehrnent rneehantsrn requtrernents and Flgure 4 1 shows the attaehrnentrneehantsrn 39Atzachment mechamsm and resmcmm Envelope Fxgure 4 51 1qu m1 Armlgunzm mnm zdmbewcmchmx 18cm Th2 1111111151311 51 1 1 Shaw 111 121112 wnh 11s 1112mm xemwved maimg 11 mtenax mm mm 5111 11111 warm 11111111111111511111111 1mm ompommplaumm 1ng 1 1 2111115 1 1 35111111111 1mm 1mm andTnnyespcmely Tray lsfnnhzslfmmthz ASAPadapm ngmjuz mmme ngmju mm71192 52 Mass Budget The mass budget shown in Table 521 includes a factor of 15 on all of the components that are not being bought as nished products Mamas 045 2 09 I HYStSfSiS I 025 I 4 I 1 I I Ltaundwtveh r39J i 3quot I ASAPAdaptor 03909 1 058635 IStrytiiures I J Bankas 1 1 1 Coatings ac 05 1 05 I Msc I Gables 1 1 15 MSC 25 1 375 Maxinrum Allowed 53 Mass Moments of Ineltia The mass moments of inertia are shown in Table 531 The values in Table 531 are well under 20 kg ml The spreadsheet used to calculate these values is given in Appendix 53 39 Ma sslkllzamant sof Iherf r MOIX39X39 kg m2 MOIy39y39 kg m2 MOIz39z39 kg m2 0287803717 0310419601 0483976707 6 CommunicationsPayload traeked anrrnals satellrte and ground stauon ltrs requrredthatthrs system be able to orbrtal und t In tlus ease L tlny txansrnltters attaehedto the txaeked anlrnals ls aformldable ehallenge Thls ls due to rnueh as posslble Figure 51 GPS Cullar Figure 52 WildCAT P115 Posmomng System slgnals speclfymg therr enaetloeauon Anumber ofmanufacturers sueh as Lotek and VECTRON39IC Aerospaee produee offrthershelf eollars for just thrs purpose In order to transrnrt tlus data to SATZ however these eollars wlll rnterfaee th devlce ls arneasly l 1 Watts A exlble quanerrwave dlpole antenna wlll be used for d 39 ent Its length wlll be 18 o ern one quarterthe wavelength forthe transrnrt frequency 402 MHz ease The resulung Effectlve lsotropreally Radated Power EIRP ls 2 514 dB In order for SAtho reeewe th5 ant slgnal 2 2xlo39 W antenna gan andreeewer 1n tlus ease mlr lM and eost Thls ls due to the ught no pun rntended and unavoldable restrretrons othe tt deadln the water Cost on the otherhand may be negotrable The ehosen antennatype 10 is a halfwave dipole of length 358 cm This length is the mean half wavelength for the various frequencies used For transmission down to the collars the frequency corresponding to this antenna length 419 MHz is chosen hence maximizing the antenna gain Using this frequency will require the least amount of power to be supplied by the satellite Minimizing power consumption is critical for this case since this transmission puts the maximum requirement on SATz s power system necessitating 1 Watt of RF power The signal will be received by the collars with a power margin of 19 Watts at the satellite s maximum broadcast distance The linkbudget for uplink and downlink between the animals and SAT2 is shown below otal Loss 2219E1 1897E1 1000E1 1000E1 data rate BER EbNo The other phase of the SAT2 mission will then be to relay the gathered information to the Purdue Ground Station PGS For this case uplink transmission power is not an issue since the ground station will transmit with 50 Watts of power and a high 14 dB gain Furthermore given the high receiving gain of the PGS antenna it is possible to minimize the transmit power of the satellite to 05 Watts As can be seen here these communications are not the limiting case The linkbudget for communications between SAT2 and the Purdue Ground Station is shown below otal Loss 1363E1 1363E1 1000E1 1000E1 data rate BER EbNo In order to simplify the complexity of the communications subsystem and to conserve power the antenna s onboard the satellite will be used for transmission and receiving of communications between both the animals and the ground station Although this will complicate the communications schedule the advantage of reduced mass and size out weighs the disadvantages In order for this scheme to work all of the broadcast frequencies must be within the bandwidth of the antenna The typical bandwidth of a dipole antenna is 10 The frequencies of 402 419 and 435 MHz used abide with this constraint Another constraint over the system is the polarization of the signal Because the Purdue Ground Station is designed to receive and transmit circularly polarized signals as is typical with AMSAT systems SAT2 must also transmit and receive in this manor In order to achieve circular polarization using dipole antennas two identical antennas must be aligned 90 degrees out of phase As a result two antennas are required to produce the correct signal SAT2 will incorporate 4 antennas in order to account for pitching of the satellite and supply some redundancy to the system One antenna will be placed on each of the four smaller sides parallel to the radius of the earth The placement of the antennas was coordinated with the attitude control system in order to account for a ip at the Earth s magnetic poles which is discussed in further detail in Section 7 The exact layout of the antennas with respect to the rest of the structure can be seen in Figure 131 Antennas Recelver Computer Modem gt gt Transmitter 4 Figure 63 Com System Layout v Component 4 i W I V Specs amp Characteristics Receiver Sensitivity 150 dBW Up to 144 Kbps Power Consumption 77 mW Size 83 X 72 X 28 mm Mass 198 gm Operating Temp 10 C to 60 C Transmitter TX437 07 W RF Output Power Up to 384 Kbps Size 94 X 72 X 28 mm Mass 300 gm Operating Temp 10 C to 60 C Modem MOD96 GMSK Modulation 10395 Bit Error Rate 2400 to 34800 bps Size 77 X 70 mm Mass 60 gm Operating Temp 20 C to 70 C All of the onboard components comprising the communications system are supplied by SpaceQuest Ltd By doing so compatibility problems can be avoided making it easier to integrate the system Furthermore all of the products chosen are already space quali ed From the antennas the signal will be routed through a RX450 receiver This receiver was chosen for its high sensitivity and frequency range The sensitivity is the minimum power of the signal that can be received in our case 7150 dBW NeXt the signal will be demodulated using the MOD96 which has error correction capabilities for bit error rates BER as low as 105 Due to the relatively low data rates at which data will be transferred the energy per bit to power spectral density of noise ratio EbNo for the received signal has a considerably large margin over that for a BER of 10395 and is therefore more than adequate In our case power margin was the far more limiting case This can be seen in the link budgets seen in Tables 61 and 62 Having been converted from a phase modulated signal to a amplitude modulated signal by the modem the data can now be manipulated and stored by the satellite s onboard computer system A similar routine is conducted in the opposite order for the transmission of data The data is mapped to an analog signal by the modem and then sent to the TX437 Transmitter which supplies an output power of 05 to amplify the signal The signal is broadcast using the same four dipole antennas The block diagram in Figure 63 visually demonstrates the integration of these components whose specifications and characteristics are listed in Tab 3 us mpmm a um um m campumm 5mm msbased mum m ssmn xeqmmems um 5mm cast and am 51mins have mt yd been candvcmd A sigm cam xeducunn m m cast am swears cmddbe realized wnh m dnmtmn afsmlar mam rm 5 wdlmg snpphex Based an m shave lmkbudgeL m PET39s law RF paw mnth mans um m swears mustbe nmmmh m m 5mm elm mn my gnm um 45 dzgnes whmh translates mm an area afmnghlyi xl km rmm whmh ammal am canbe received m anyanz m s Figure 5 4mm 1mm appmxxmmelyi mums funk spacecm m ywenhls ma Emunmvack 55m 54 mpmmmumm gaommy m Eanh39ssnrface 5 pumanzd m masafthls 92 this mmmm three mmmzs duh xsalhh mm PETsmthz mamdthz mummyquot mexchmge axxmm m numbemfammals mummy canmtzdmanyaf m gmnanas amad 39lszxme DMsmn Mum Asses TDMA schzme wdlbe yzd Fm asthz swears camsmtnvlzw m callus w mmnhmdmm a dumpquot s mtmemamdan y m whim m mxt e psses acc equmxa ammmnm a z psses adayls g teed un Faxhlghlemmldzsthz number pa 5 pxdaymcxe a afeach mm wmcmammm s n 15 cleanlmfanhls schzmz mm sucessf 1nscnncal um cammunlcaonns k2 aceanhzkasslgmdumzs Hawzvex thls pmb smce scannpime aw slacks mm m GPs receiver m m mum am a dayfmm gmnnd mm 6 Ob39ts 28 its 2b39ts 6bis1 24 its 00 its Legend I Guardl CBR Bowl 3 Data GUardl Burst from animal Guard Guard time to avoid transmissron overlap CBR 1 second cw ID Animal ID number Data Animal e2ncode GPS data Complete frame Acknldg SAT received data from these IDs Ephem Updated SAT2 ephemeris 3 minutes Note Frame for ground station 6 Ob39ts 28 its 2b39ts 6bis4 74 its 00 its mm 39 at39 S39mquotart these Burst from SATZ 254 second max Figure 65 Burst and frame structure in the modi ed TDMA scheme employed for animal spacecraft communications Based on the burst and frame description in Figure 65 a maximum of 230 animals can be located in each 3000x3000 km area There are roughly 211 blocks of 3000x3000 km in the Earth s surface but since there is no guarantee that a wild animal will stay in its assigned area a smaller number of larger size areas 55 has been de ned around the world based on natural obstacles such as mountain ranges coastlines etc that would limit the travel of a land animal For marine animals such as whales which have large travel areas an even lower limit on the number tracked may have to be imposed For this communications scheme to work it is helpful that the orbit is sunsynchronous because the local time at which the PTT will transmit will can be made to vary only according to the latitude at which the animal is within its prescribed area instead of requiring a more complicated logic To communicate with the groundstation a data rate of 14400 bps will be used in order to accommodate all the animal data from 1 day within 1 groundpass It will take 155 minutes to complete the uplinkdownlink sequence with the groundstation During the time that the spacecraft is communicating with the groundstation it cannot communicate with any animal This reduces the number of available passes for animal communication in North America to 45 per day In addition it will be necessary to upgrade the Purdue Ground Station since at this time it is only capable of 400 bps 7 Attitude Determination and Control System ADCS 71 Major Requirements Satellites must be oriented properly in order to ful ll the mission statement properly For SATZ the antennas must be nadir pointing with a pointing error of less than 25 degrees in order to minimize communication losses The attitude determination and control subsystem takes this as well as the solar panel orientation into consideration in order to make an effective design 72 Disturbance Environment The space environment in Earth s orbit offers several external torques that ADCS must control or tolerate These torques include a gravity gradient a solar radiation pressure a magnetic field and aerodynamic forces Detailed calculations of each disturbance are included in Appendix 7 and the results are shown below in Table 721 T of Disturbance T 0 86087E4 24603E8 As seen from Table 721 when comparing cases the worst magnetic field disturbance has the highest magnitude about 3 orders of magnitude greater than the gravity gradient the solar pressure and any aerodynamic torques 73 Selection of Control System Because SAT2 does not require a pointing accuracy and its time life is infinite a passive control system was selected During the early phase of the design process two systems were studied and one was selected Both system are listed below 731 Gravity Gradient Boom The Gravity Gradient boom uses the inertial properties of a satellite to keep itself pointed toward the Earth The idea of a gravity gradient boom can easily be explained when the satellite is modeled as a dumbbell The weight that is closer to the Earth would be pulled more than the weight farther from the Earth The satellite is not moving closer to the earth which causes the dumbbell to rotate until it is vertical Advantages of gravity gradient boom apart from being a passive system are that it allows for earth pointing However satellites with gravity gradient booms tend to nutate A nutation damper is required to counteract this motion Another disadvantage of the boom is that it involves a deployable mechanism which can be complicated and expensive With the advantages and disadvantages in mind a stability analysis was performed Details about the boom and the calculations are included in Appendix 7 Below is the list of the size of the boom required to keep the pointing error within 20 degrees Stowed volume 102 X 115 X 264 mm Deployed volume 102 X 115 X 6264 mm Mass 22 kgI 732 Passive Magnetic Permanenm magnet mounted in the satellite are used for passive stabilization and control They are used in order to make attitude determination a more reasonable task by holding the satellite in a mathematically predictable orientation With appropriate dipole strength and orientation the magnem can be used to keep the satellite oriented with the local Earth magnetic field vector Figure 732 shows the motion of SAT2 over a period of quarter orbit i Figure 732 Motion of SAT2 for a quarter orbi The motion shown in Figure 732 has been called a controlled tumble where the satellite ips over near the NorthSouth pole in order to align its magnetic dipole 39 the Earths magnetic field The satellite will ip twice per orbit near the poles The sizes of magnem required depend on the magnetic strength needed and tend to be relatively small The permanent magnet will not normally make the satellite nutate or spin Hysteresis rods will be used in order to be certain that the satellite will not spin out of control Permanent magnem were chosen over the gravity gradient boom mainly because of its compact size and ease of design integration 74 Component Selection and Sizing 741 Permanent Magnet Two different types of magnets were considered Alnico and Neodymium Iron Boron Alnico has been used on several small student built satellites such as Webersat Alnico is relatively inexpensive and has medium to high strength and very high temperature stability but low resistance to demagnetization Since the Earths magnetic eld is relatively weak when compared to satellite magnets the concern for demagnetization can be neglected Neodymium Iron Boron NdFeB is known as rare earth magnets because it composed of materials from rare earth group elements NdFeB has a very high strength very high resistance to demagnetization but low temperature stability Since high temperature is not a case for SATZ NdFeB was selected Base on the database collected the size and strength of each magnet was chosen to be Size 075 x 075 x 06 inch Magnet moment 26E3 EMU 742 Hysteresis Rods Hysteresis rods are used to keep the spin from increasing The idea was based on AMSAT satellites which have used hysteresis rods to control the spin rate Hysteresis rods are composed of 49 percent hyperm steel and are hydrogen annealed These rods generate eddy currents when passing through a magnetic eld When the iron rods move through the Earths magnetic eld an electromotive force emf is produced The calculations involving hysteresis rods are located in Appendix 71 The sizes ofthe rods were based on the design of Spartnik by San Jose State University and are listed below Number of rods 4 Length 14 inch Diameter 014125 inch The hysteresis rods will be place along the plane perpendicular to the spin axis Z axis 75 Predicted Performance Moments of inertia of SAT2 that were used in the simulation were calculated to be Ixx 02878 kgmz Iyy 03104 kgmz IZZ 04840 kg mz Matlab was used to simulate the performance of SAT2 under the in uence of the Earths magnetic eld The main goal of the simulation is to determine if the passive attitude control system will perform as planned Several matlab scripts were written in order to simulate the control system and the earth magnetic eld was model after the International Geomagnetic Reference Field IGRF Calculations can be found in Appendix 7 18 SAT2 is expected to ip twice in an orbit with respect to the Earth therefore it is expected to rotate one revolution with respect to the rotating frame Complete explanation of the frames are located in Appendix 7 The simulation is restricted to a quarter of an orbit because a singularity will occur at that point The test case was conducted without permanent magnets no torques on board and all the initial conditions set to zero The result confirmed that SAT2 will remain pointed throughout the simulation The pitch angle appears to be changing because the rotational frame rotates with respect to the Earth s fixed frame The second simulation Figure 72 shows the motions of SAT2 with 2 permanent magnets on board and zeros initial conditions The roll and yaw angle appears to be zero throughout the motion the peak at the end occurs because singularity that occurred For a quarter of an orbit SAT2 appears to be pitching 90 degrees as expected and the pointing error appear to be within 20 degree when ignoring the singularity effected Roll Pitch n n 20 A 50 S 3 3 4 3 V m 39 quot 60 i 100 2 80 4An 4nn 0 10 20 30 0 10 20 30 Time min Time min Yaw Pointing error n an 3 a a 50 E 10 w m B E e E 100 t o a E O 4quot 4n 0 10 20 30 0 10 20 30 Time min Time min Figure 72 SATz motion with 2 permanent magnets on board By increasing the initial spin rate of the satellite to 01 reVsec the motion of SAT2 appears to be similar to those without magnets With these results it appears that the permanent magnets on board should be able to perform as designed Since SAT2 communication links will be preprogrammed there will be no attitude determination system required 8 CampDH The purpose of the CampDH subsystem is to provide overall control of the spacecraft subsystems and to collect and store telemetry and customer data for transmission to the groundstation Based on the concept of operations discussed in Section 2 of this report the maximum amount of user data that the CDampH data is required to store per day is 1606 Kbytes The total available memory is 4 Mb which allows for the storage of up to 2 days worth of maximum data before transmission Since there are 56 passes with the required groundstation visibility see Section 3 it is unlikely that this capacity will have to be used Even in the event that it is decided to increase the size of the available memory at a later date this can be done with a minimal impact on mass moments of inertia size and power consumption up to 16 Mbyte In addition to storing collar data the CampDH subsystem must carry out the following functions propagate its orbital ephemeris to determine groundstation passes collect and process telemetry information exectute ground commands and control the power subsystem Figure 81 is a schematic of the CampDH subsystem and its relation to the other spacecraft subsystems Telemetry Board 3 Modem FIIghtComputer I I I Trans mitter t Memory Board Figure 81 CampDH subsystem schematic The components of the CampDH subsystem their manufacturers and their principal characteristics are listed in Table 81 More details on these components can be found in Section XIV of the Appendix In selecting the ight computer and modem other options significantly cheaper were studied but the SpaceQuest products were selected because they are space qualified they are radiation hardened and they are designed to work together This means they will require little or no modification radiation testing might be reduced and integrating them with the rest of the spacecraft components most of them also from SpaceQuest will require less effort 20 Modelamp f Characteristics Component Flight Computer SpaceQuest Inc Model FCV53 5 MHz l Mbyte EDAC RAM Memory Board SpaceQuest Inc Model FMl6 3 Mbytes EDAC RAM 10 mW at idle Modem SpaceQuest Inc Model MOD Fire code to reset FCV53 96 on ground command Due to this feature considered part of both communications and CampDH Telemetry Board TBD Assumed 01 W Probably based on Tattletale consumption Model 8V2 data logger or Probably can manufacture similar at Purdue See requirements below A very important aspect of having electronics in space is their sensitivity to radiation effects The degradation of electronic components coming from the Total Ionization Dose TID can be reduced by aluminum shielding except high energy protons by limiting the amount of time in orbit or by selecting an orbit with a benign environment For SATZ the aluminum trays used to hold the electronic boxes have been arranged in a way that they also act as aluminum shields for the most critical components the ight computer and memory board see Figure 5111 in Section 5 of this report In addition all electronic components are housed within aluminum boxes SAT2 is designed for an operational life of one year so this should help limit the TID Unfortunately the high inclination orbit selected takes SAT2 outside the protection of the Earth s magnetic field when ying over the poles and through the South Atlantic Anomaly Based on previous satellite designs and the short lifetime it is expected that no additional shielding beyond that described above will be required but further work will need to be done in this area The other radiation related area needing consideration is Single Event Effects SEE caused by high energy protons and cosmic rays Based on reports from other missions SDRAM and processors like those selected for SAT2 tend to be relatively immune to Single Event Latchup SEL leaving only the less damaging Single Event Upsets to be dealt with The ight computer and memory board selected deal with SEU through the use of Error Detection and Correction algorithms A total of 4 Mb EDAC memory is available through the use of 12 Mb of physical memory The main criteria for the selection of the above mentioned components was their ability to meet the requirements for processor speed and memory size determined from the analysis detailed in the following table 21 Required Ephemeris Propagation Command Processing Telemetry Processing Power Management Operating System Local Executive n Runtime Kernel lO Handlers m BIT and Diagnostics Margin calculations Requirement Uncertanties b 50100 411129 50100 The telemetry acquisition part of the CampDH subsystem has not been fully de ned yet since based on other university small satellite missions this is usually a custom built component tailored to the particular requirements of each spacecraft The following table details the inputs that up to now are required from the telemetry board level level level temperatures as a attitude determination method status Based on the above requirements a fairly typical arrangement is envisioned for the telemetry board similar to those described in the literature for other university run missions It is possible that such a data acquisition system could be easily developed by students at Purdue using components from a variety of suppliers It should be noted that at this time the telemetry board is expected to collect data directly from the required 22 components and only for the purpose of transmission to the ground but it is possible that some devices that collect the same analog ie temperatures and voltages data for control purposes such as the battery charging unit can provide part of the data to the telemetry board or they can both use the same collection units Since the spacecraft is relatively simple only a limited number of commands will be required to control it Most of the spacecraft control during normal operations ie when to charge batteries when to switch to receive transmit mode etc will be carried out by the onboard computer leaving only commands for the afterlaunch activation and restarting after an emergency switch to safe mode required from the ground 23 9 Power Computer 9425 V Regulator 33 V Modem V 1 i 1 Memory Board Solar Array Telemetry Battery Board Char in Regulator 5V Receiver Unglt g Battery i 1 Transmitter Shunt Dump Figure 91 Schematic of the power subsystem The power subsystem must be able to provide power to operate all satellite components required for normal operations at all times In the event that the primary source fails the secondary source must be able to provide power for enough time for ground controllers to troubleshoot the problem and attempt corrective action Figure 91 above is a schematic of the power subsystem There are four operational modes depending on whether the transmitter and receiver are operating at a given time and at what power level they are operating Essential components are operating constantly such as the CampDH and power subsystems along with the receiver The following table lists the different operating modes In Safe it listens for The power subsystem uses GaAsGe solar panels as the primary power source and Liion batteries for eclipseemergency operation The excess power from the solar array is dissipated by a shunt regulator depending on the load on the power subsystem The battery has a charging unit that follows the battery manufacturer s prescribed charging procedure as shown in Figure 92 24 1O Charge Delivered 0 Cell 5 Voltage 3 CA Charge Rate O 2 Time hrs Current Limit Constant Voltage Figure 92 Battery charging pro le The loads are arranged in two groups according to their required input voltages and each of these groups is controlled by regulator that provides the required voltage as shown in Figure 91 The following table lists the power subsystem components their manufacturers and their principal characteristics A more detailed description of the components is available in Section 9 of the Appendix Ta Component Mode amp f Characteristics Solar Array Spectrolab Dual Junction 120 7x312 cm cells total GaAsGe 1821918 area of026 m2 efficiency X cells in series y in parallel Battery Valence Inc Model 593 3 cells in series to provide VC315590 111 V and 1432 Whr 20 Cto60 C operating range Battery TBD 01 W consumption Charging Unit Based on data from other satellites and National Semiconductor database Regulator SpaceQuest Inc Model VC9 Input Voltage between 8 and 28 V Output voltage between 27 and 16 V maintained to i1 15 loss assumed Shunt Dump TBD Probably able to produce it at Purdue The power subsystem was designed based on the power budget shown in Table 93 which details the power required in each operating mode including a 30 margin for 25 line losses and unforeseen requirements In preparing this power budget the transmitter was assumed to be 20 efficient and the receiver was assumed to draw 0125 W while receiving based on a similar receiver by the same manufacturer All other values are from the manufacturer specifications Based on the above power budget the battery was sized and from the combined requirements of operations at the highest power when transmitting to the animal collars and the battery charging the solar array was sized To determine the size of the battery the characteristics of the selected cell shown in Figure 93 were used This cell was chosen was picked due to its small size both in terms of physical size and capacity its high energy density and the fact that similar cells have been used in space before 26 IBM Hall Ime r v E 2 m a crw n 39n 1 M m m 1 Witr quotn m u m m an m u w HEW mum EM LII Charmu39ll m gt i39 Figure 93 Cell characteristics In sizing the battery it was assumed that the groundstation communication and as many contacts with animal collars as time permits take place during the worst case eclipse which occurs when the normal eclipse is followed by a Moon eclipse as discussed in Section 3 of this report This results in a very low depth of discharge of 12 but by sizing the solar arrays large enough to fully charge the battery from zero see Figure 92 it provides the full battery capacity for emergency operations This feature gives approximately 14 hours and an average of 4 groundstation passes for ground controllers to attempt to resolve the problem and still have the ability to fully recharge the battery within a maximum of 12 orbits and resume normal operations To achieve this the battery was sized to provide discharge at less than C5 C is the battery capacity see Figure 93 This also allows the battery to provide the required capacity for normal operations up to 10 0C As mentioned before the solar array was sized to be able to provide enough power to charge the battery initially at C2 and simultaneously operate at the highest power consumption mode even though in normal operations the usual rate will be around C20 due to the high battery voltage caused in turn by the low DOD In sizing the array it was assumed that only the two 40x40 cm satellite faces are covered by solar panels and the average pointing angle to the Sun is 45 degrees based on the attitude history analysis in Section 12 of this report Since one of this faces has to carry the payload adapter for the launch vehicle and this does not leave enough area for the solar arrays it was decided to place solar arrays also in the four 40x18 cm faces Table 94 shows the power margins at all operating modes More details on the calculations performed to design the power subsystem can be found in Section 9 of the Appendix to animal station animal GS assume margin will higher 27 Thermal Control 101 Objective and Requirements The purpose of the thermal control subsystem is to keep the temperature of the spacecraft at a temperature that allows operation of all the spacecraft components particularly electronic ones The following Table shows the operating and storage temperatures of all the spacecraft components Based on this the allowable temperature range for the spacecraft is between 710 and 60 0C 102 Major Heating Sources The average length of time spent by the spacecraft in sunlight is 3956 secondsThe major sources of heat during this time period are 1 Solar ux 1418 Wm2 2 Albedo effects 3035 of solar ux 3 Heat generated by instruments on the satellite Max 8303 W Min 0808 W 4 Earth Infrared Radiation 237i21 wm2 The average length of time spent by the spacecraft in eclipse is 2092 seconds The only source of heat during this time period is the heat generated by electronic components onboard the spacecraft 103 Predicted Spacecraft Temperatures To predict the spacecraft temperature experienced by the spacecraft in orbit a two stage analysis was performed 1031 Steady state approximation or lumped capacitance method Assuming the satellite is a point in space and is in steady state heat transfer the equilibrium temperatures of the spacecraft are calculated These temperatures represent worst case scenarios Due to its low orbit the spacecraft is not going to 28 spend enough time in either eclipse or sunlight to reach these temperatures These equilibrium temperatures are estimated by using the conservation of energy equation Qahrnrhed Qampm 39med 0 Thus by inserting the values for solar ux albedo 1R emissivity emissivity and absorbtivity for solar panels since they cover most of the satellite the equilibrium temperatures are predictedto be mesunlight 42 DC Tmmeclipse 90 DC Refer appendix for the spreadsheet of the analysis To obtain these temperatures the following assumptions were made own I J m rnrrrtrof asphere o The emissivity of satellite is equal to that of solar panels 1032 Transient Thermal Analysis The orbital period of the satellite is small and it sees sunlight and eclipses fast and hence it never actually sees steady state but is always under transient effects For the transient temperature analysis the following differential equation for the isothermal spacecra in sunlight was numerically integrated Wit T aas 6 e ArrT4 it where m is mass of satellite c is specific heat of the satellite at is the absorbtivity e is the emissivity and u is the StefanBoltzman constant The model used includes the effect of initial temperature within the launcher fairing between 10 and 25 DC as speci ed by the primary payload and the effects of aerothermal heating a er fairing separation by adding an average heat ux Figure 101 shows the time history of spacecra temperature after fairing separation Temperature 60 Ovbtls Fzgare 101a Spacecraft surface temperatare me hzstory for 15 mm Mum pr H I w I Orbits Figure 101b Spacecraft surface temperature time history for 300 orbits Based in the above Figures and calculations for up to 3000 orbits the maximum temperature experienced by the spacecraft is approximately 35 0C and the minimum temperature experienced is approximately 712 0C The maximum temperature is 25 0C below the maximum allowable operating temperature and the minimum is 2 0C below the minimum allowable operating temperature However this difference is small and can be easily overcome by the addition of insulation such as MLI if further analysis justifies it Actual data from similar spacecraft currently in orbit PCSAT suggests that the temperature range will be approximately between 78 and 15 0C The following assumptions were made in the transient analysis The satellite is completely covered with solar cell 00085 6 0825 Cp900jkgllt Aluminum The satellite is injected at terminator into the dark side 104 Thermal subsystem components The thermal control subsystem is currently designed as a completely passive system Typical passive thermal control subsystem components are 1 Insulation 2 Coatings 3 Isolators Thermal coatings having similar 6 0 ratios to that of solar panels will be used on the four 40 x 18 cm smaller sides if they are not covered by solar panels Alternatively coatings with a lower ratio that will help increase the minimum temperature might be used 30 To prevent heat transfer between components due to conductance thermal isolators lowconductive materials like NARMO berglass Stainless Steel and Titanium can be used 105 Future Work The temperature predictions are based on the assumption that the satellite is a sphere Temperature based on the actual geometry of the satellite need to be obtained The indiVidual r of 39 39 r on the satellite need to be found For this radiative coupling analysis and a more sophisticated approach such as Finite Element Analysis will be required The temperatures of the satellite after a few months which might change due to thermal degradation of thermal coating and that of solar panels These effects will also require analysis 31 1 1 Structures and Mechanisms 111 Major Requirements SAT2 was designed for launch aboard the ASAP Ariane Space Auxiliary Payload micro satellite launch platform ASAP allows a maximum mass of 120 kg and a maximum volume during launch of 60cm x 60cm x 71cm The maximum allowed mass moment of inertia about any axis is 20 kg ml The maximum axial and lateral forces as well as the axial and lateral frequencies are given below in Table 111 1 which also shows a summary of the ASAP requirements SAT2 Group members imposed several additional requirements The structure was designed to use only off the shelf aluminum parts The issue of machining smaller sized square channels in order to reduce weight was addressed but the weight savings were not great enough to warrant the extra time and cost expense needed for custom machining shapirosupplycom and aircraftsprucecom were the suppliers used to compile an Aluminum stock database Appendix 11 l for parts determination and ordering Also a maximum launch volume of 40cm x 40cm x 40cm a mass less than 50 kg was imposed 112 Initial Design Assumptions and Structural Concept Square channels were picked for the primary structure to ensure a robust foundation 5 x 5 x 058 was the smallest size available which by intuition was very big compared to the selfimposed limit of 40 cm3 This size square channel was determined to be oversized compared to the overall size of our structure However we had already agreed to use only off the shelf parts We then knew that our primary structure would be more than adequate to support the ASAP required loads Because of this the details of the structural design of SAT2 were then assumed to be based on meeting the frequency qualifications as with most small satellites 113 Material Selection The structure is composed of 606lT6 AL This is atypical satellite material and was chosen because of the low loads on the spacecraft availability and most importantly because it s the easiest to machine 2024T3 AL has a higher yield and ultimate load capability but is slightly harder to machine and is more expensive per sheet Appendix lll shows a cost comparison 32 114 How Spacecraft Supports Concept of Operations By choosing off the shelf stock SAT2 will be easy to construct as a student built satellite By having no moving parts cost is reduced and design time signi cantly reduced All stock was chosen to be easy to use readily available and cost effective 115 Vibration Analysis In order to determine the frequency the structure was modeled in QED as two separate components A primary frame structure and single tray were modeled separately When the frame was analyzed alone See Figure 1151 the lowest frequency occurred in the 1st mode ofthe frame and was 136 Hz See Table 1151 Figure115139 Snapshot afframe vibration sequence The trays were modeled separately as 40cm X 1163cm X 016cm sheets with one side free initially The frequency for the 1st mode was 163234 Hz Then the tray was modeled with all sides fixed The frequency of the 1St mode was found to be 842 Hz It was decided that the trays would then be fixed on all 4 sides even though some weight would be added Values for the first few modes of both the primary frame structure and the tray are given in Table 1151 39ri fer Frame Structure Tray one side free Tray All sides fixed Mode Frequency Hz Frequency Hz Frequency Hz 1 13644 163234 842231 2 13644 252488 891918 3 162167 391212 981664 4 188017 578542 111809 5 195818 814809 130594 6 219566 836558 154752 7 252393 930566 184338 8 259925 107744 2193 9 259925 110097 226265 10 283915 127072 231681 33 The assumption was then made that a combination of these two models would bring down the tray frequency some closer to the value of the primary structure giving an overall structure frequency of at least 13644 Hz Thus it was assumed that the value of the combined primary structure and tray frequencies would never dip below the value for the primary structure This analysis is not very accurate due to the very simplified nature of the assumptions 116 Structural Analysis 1161 Critical Fastener Analysis The load analysis was reduced to back of the envelope calculations to show that the fasteners that transferred the load from the Ariane 5 to SAT2 would be well under their yield strength The approach that was used assumed a overly large safety factor of 20 in order to account for the grossly primitive analysis This approach gives a spacecraft that is unnecessarily overweight This safety factor was added to the maximum lateral load of 79 G s in order to determine the largest force that the Ariane 5 would impart on the SAT2 structure A total force of 3250 N was determined This total force was then divided between the total number of fasteners that took the load that would be transferred from the Ariane to SAT2 and those values were compared to yield values for various fasteners See Appendix 11611 for calculations The load that each fastener needed to withstand 92lb in shear per fastener was well under the yield value of several fasteners that were given in Mil 5 The other areas that were potential places for failure were the 4 fasteners that held each tray to the primary structure A special fastener assembly was designed as a way to attach the tray square channels to the exterior frame See Appendix 11612 for schematic Most 0125 D bolts that would attach this special assembly have adequate shear and tensile strength to withstand the shear load and or tensile load that would be imparted upon them 1162 Tray De ection Analysis The de ection of the sheets was also investigated The panels were divided into three sections each fixed on all sides These sections of roughly 40cm x 11cm x 016cm were analyzed with an inertial load equal to the largest tray supported mass with a safety factor of 2 The de ections are shown in Figure 1162 The maximum de ection was found to be 299e1 inches This de ection neglects the fact that the beams that support the trays are going to de ect some as well 34 Figure 1162 Plat afmwcimum de ection in inches 00299 aftray with all edges xed 117 Fuither Analysis Needed The next step in the structures design would be a detailed inertial load and Vibration analysis in ANSYS or ABAQUS This was not completed for this class due to time constraints Also a complete set of detailed engineering drawings is necessary 35 l Summary of Spacecraft Requirements and Capabalities The following table describes the most requirements the spacecraft must meet and how the current design meets them Table 121 Spacraft J vs capabilities table requirements both mechanical and electrical Only required electrical connection is for battery trickle charging as required Requirement Capability Comments 1 Ability to track a large Can track 12650 animals trough number of animals PTT UHF link and modi ed TDMA scheme 2 Worldwide coverage 800 km sunsynchronous orbit 3 Meet launch vehicle See Table 122 below Based on the preliminary analysis so far performed 4 Must use the Purdue Ground Station UHF transmitter and receiver 1134 dBW power margin 138 dB EbNo margin The PGS will have to be upgraded so it can receive at less than 20 0 pointing error except when ipping at the poles 339 dB uplink 139 dB downlink 14400 bps EbNo margin 6 Must use existing GPS Lotek Vectronic collars WildCat collar and PTT PTT 35 dBW uplink 28 dBW downlink power margin 133 dB uplink 126 dB downlink EbNo margin 7 Nadir pointing Passive magetic stabilization with Pole ipping does not interrupt communications due to the omni antenna a1rangement in the satellite 8 Avoid spinning Histerisys rods damp out spin Toreduce power loss from the solar array 9 Must provide enough memory for operations 3614 Kbytes required 4000 Kbytes EDAC RAM provided radiation environment 10 Processor must be fast 5 MHz available enough for the required throughput 028 MHz required 11 Must survive the space EDAC RAM Aluminum tray arrangement doubles as shielding 36 All electronics housed in aluminum cases 12 Must provide enough 8707 W from the battery max Also provides the power for normal and 155 W from solar array required voltages emergency operations Smallest margin is 183 during for operation of all emergency operations components 13 The spacecraft must Transient analysis indicates The minimum stay within the required temperatures between 7 12 and 35 temperature can temperature range 7 10 0C raised by the use of to 60 0C coatings insulation 14 Structure must be Simple low cost Aluminum 3 J at Purdue Structure 15 Structure must carry all All components housed in loads and hold all aluminum trays held by an r aluminum frame Table 122 Launch vehicle requirements and current design comparison EQUIRED Current estimate Volume max 600mm X 600mm X 710mm 400mm X 400mm X 180mm Weight lt 120 Kg 2078 lXX lt 20 mquot2 0302 lyy lt 20 mquot3 0329 lzz lt 20 mquot3 05196 omega aXial min of 90 Hz 135 omega lateral min of 45 Hz 135 FaXialstatio 75 g39s 1185 Flateralstatio 6g39s 9 Cost Estimate The following table is a conservative estimate of what the spacecraft is expected to cost It assumes that most work will be carried out by students but they will be supervised by two hired engineers The 250hour cost for each of the engineers includes benefits and administrative costs Table 123 cost estimate T Generator 1 750000 for 1500 hrs Future Work The orbit must be reanalyzed based on the actual orbit of the primary payload once a launch is obtained Once a launch opportunity arises the spacecraft must be redesigned around whatever launch vehicle it is with a full static and dynamic structural analysis The communications scheduleprocedure must be analyzed in more detail with help from someone knowledgeable in animal migration patterns The accuracy of the spacecraft and PTT clocks must be determined in order to reduce the large guard times around each PTT burst A more detailed analysis of the losses in the link budgets must be carried out A full analysis of the integrated passive magnethisterisys rods must be carried out including a more detailed of the Earth s magnetic eld The telemetry board must be fully designed to meet its requirements The interfaces between the various electronic components must be fully analyzed The onboard and ground commands must be fully specified and their effect on software development assessed The software functions both for the Flight Computer and the telemetry board must be fully defined A more detailed analysis of the power produced by the solar panels based on the attitude history must be carried out A more detailed thermal analysis of the spacecraft which predicts also and gradients within the spacecraft must r r be carried out A more detailed analysis of the structure must be carried out An analysis of electromagnetic compatibilityinterference must be carried out An assessment of testing requirements capabilities methods and schedules must be carried out An assessment of what facilities such as clean rooms testing facilities need to be built at Purdue must be carried out 38
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