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YP in space2018_bcn_all_slides_theory_x2

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Theory Sessions of the IEEE Young Professionals in Space boot camp - UPCBarcelonaTech July 2018

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YP in space2018_bcn_all_slides_theory_x2

  1. 1. 1 IEEE YPinSpace 2018 © UPC‐BarcelonaTech ‐ 2018 1 IEEE Young Professionals in Space … about this bootcamp IEEE YPinSpace 2018 © UPC‐BarcelonaTech ‐ 2018 2 Organization Morning: Invited Lectures (see web site) Theory and Practical Sessions Alternate: Theory Hands on session 1T Orbits 1P Orbits Lab using the Princeton Satellite Toolbox and  Orbitron 2T Space Environment and Thermal Control 2P Thermal Control Analysis using the Princeton  Satellite Toolbox 3T On Board Data Handling System (OBDH) 3P OBDH: Getting stated with Arduino 4T Attitude Determination and Control System (ADCS) 4P ADCS: Understanding Sun sensors, reaction wheels,  magnetorquers, GPS… 5T Electrical Power Supply System (EPS) 5P EPS: Understanding solar cells, MPPT, battery chargers, batteries… 6T Telemetry, Tracking and Control System (TTC) 6P1 TTC: E2E communications and transceivers in radio  amateur bands 6P2 TTC: Ground Station: Manufacturing a Yagi antenna and using the RTL to monitor the spectrum and decode APRS packets 7P E2E tests, rocket launch, and telemetry tracking
  2. 2. 2 IEEE YPinSpace 2018 © UPC‐BarcelonaTech ‐ 2018 3 Housekeeping information Hands on sessions: Electronics Eng. Labs C4 building, floor ‐1 Theory Sessions: Agora room Plaça Telecos Lunch Unity Cafeteria Plaça Telecos Coffee breaks B3 building – basement Plaça Telecos Lunch boxes will be  provided for Saturday trip to Alcolea de Cinca Unity Cafeteria Plaça Telecos Invited lectures Registration 16/07/2018 Water rockets side event IEEE YPinSpace 2018 © UPC‐BarcelonaTech ‐ 2018 4 Satellite subsystems: • Electrical Power Supply (EPS):  solar panels, battery charger, and batteries • On Board Computer (OBC):  CPU, memory • Communications Systems (COMS) transmitter(s) + antennas • Attitude Determination and Control System (ADCS): sensors + actuators (reaction wheels & magnetorquers) • Payloads (P/L)
  3. 3. 3 IEEE YPinSpace 2018 © UPC‐BarcelonaTech ‐ 2018 5 EPS ADCS OBDH COMMS P/L SATELLITE GROUND SEGMENT RTL‐SDR (spectrum monitoring) IEEE YPinSpace 2018 © UPC‐BarcelonaTech ‐ 2018 6
  4. 4. 1 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 1 IEEE Young Professionals in Space T1. Introduction and Orbits T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 2 • Objectives of the lesson - Learn the phases and segments of a satellite mission - Learn about Orbits: • Understand Kepler laws • Understand Orbital Parameters • Understand Perturbation factors • Learn the different types of orbits and their applications • Learn how to compute different types of orbits and the associated visibility time • Learn how to compute the reentry time
  5. 5. 2 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 3 • Phases of a Space Mission (1/2) – Phase A: Feasibility study (8 to 12 months) – Phase B: Detailed Definition (12 to 18 months) – Phase C/D: Development, Manufacture,  Integration/Test (3 to 5 years) – Phase E: Launch Campaign  • Pre‐launch: prep. ignition, separation umbilical cables • Launch: ignition, continuation, separation of different  phases • Orbit transfer: launch orbit to operational orbit T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 4 • Phases of a Space Mission (2/2) – Phase F: Mission Operations • Commissioning phase: pull payload out, verification  • Mission operations: attitude, repositioning, orbital  maneuvers, fuel limitation  satellite lifetime • Decommissioning phase:  – GEO: impulse to higher altitude orbit – LEO: controlled re‐entry
  6. 6. 3 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 5 • Necessary expertises – Propulsion, launch and re‐entry – Orbital analysis, navigation, attitude control and maneuvers – Structures and materials – Thermal and radiation control – Sensor development – Telecommunications – Signal Processing – Mechanisms – Electromagnetic Compatibility – Production and energy storage – Simulation – Project management, data distribution, etc. T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 6 • A space mission consists of: – Launcher: transports satellite from ground to space – Space segment: spacecraft – Ground segment: ground station and mission control SMOS launch [credits ESA]
  7. 7. 4 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 7 • A space mission consists of: – Launcher: transports satellite from ground to space – Space segment: spacecraft – Ground segment: ground station and mission control SMOS launch [credits ESA] T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 8 • A space mission consists of: – Launcher: transports satellite from ground to space – Space segment: spacecraft – Ground segment: ground station and mission control SMOS launch [credits ESA]
  8. 8. 5 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 9 • A space mission consists of: – Launcher: transports satellite from ground to space – Space segment: spacecraft – Ground segment: ground station and mission control SMOS launch [credits ESA] T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 10 Orbit selection depends on application and payload Examples: Meteosat (Geostationary) ‐ Always looking at  the same point ‐ 3 satellites required for global coverage GPS (Medium Earth Orbit) ‐ At least 4 satellites in  view are required OKO URSS Spy Satellite (Molniya orb) ‐ Large fraction of the orbital period over the region of interest ‐ Telecom at high latitudes and spy
  9. 9. 6 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 11 Kepler’ s laws: 1.The secondary body describes an elliptical orbit around the primary body. 2. The area swept by the radio‐vector going  from the primary body to the secondary  body is the same in equal time intervals. 3. The orbital period is determined by the  average distance that separates them and by  the mass of the primary Eccentricity e Type of Orbit 0 Circular 0 < e < 1 Eliptical 1 Circular 1 < e Hyperbolic T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 12 Types of orbits: ‐ By height: ‐ LEO: Low Earth Orbit ‐ MEO: Medium Earth Orbit ‐ GEO: Geosynchronous Earth Orbit ‐ HEO: Highly Elliptical Orbit (e.g. Tundra, Molniya) ‐ Non‐geocentric orbits: interplanetary navigation
  10. 10. 7 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 13 6 Orbital Parameters (derived from the 2‐body Newton equation): ‐ Ω: longitude of the ascending node on the equatorial plane ‐ Ψ: inclination of the orbital plane with respect to the  equatorial plane ‐ γ: argument of the perigee at the ascending node ‐ a: semi‐major axis of the ellipse ‐ e: orbit eccentricity ‐ tp: time of passage at the perigee (reference initial time)  ORBIT < 90 direct = 90 polar > 90 retrograde 2 2 2 2 ˆ ˆ 0 1 ( ) 2 d r r dt r m mV GM r          2 1 cos h r e     T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 14 6 Orbital Parameters (derived from the 2‐body Newton equation): 2 2 2 2 ˆ ˆ 0 1 ( ) 2 d r r dt r m mV GM r          2 1 cos h r e     : true anomaly e: eccentricity 0  e  1: elliptic & circular orbits 3 2 1 2 sin tan tan 2 2 1 a E e T M t E e E M T e                
  11. 11. 8 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 15 Orbital period 2 , 2 84,4 , 6366 , k GM 3,986 10 m /s For a circular orbit: 2 , Ex.: TRANSIT Satellite h=1075km → v=7.3 km/s and T=106 min T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 16 Perturbations that modify orbital parameters Osculating parameters: 6 orbital parameters as a function of time • Asymmetry in the Earth’ s gravity field: Earth’ s gravitational potencial ‐ ,  : Geocentric latitude and longitude ‐ Pn:     Legendre polinomials ‐ Pn,m:  Legendre functions (1st kind)  ‐ Jn, Cn,m, Sn,m: experimentally determined coefficients          2 , , , 2 0 1 sin cos sin sin n n n n n n m n m n m n m R V J P r r R C m S m P r                                    
  12. 12. 9 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 17 Gravity Field Map (GOCE Mission) http://www.esa.int/esaLP/ESAYEK1VMOC_LPgoce_1.html#subhead2 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 18 Orbital Perturbations (i): ‐ Atmospheric losses: ‐ Due to atmospheric drag orbit circularizes ‐ For a circular orbit: ‐ Solar radiation pressure: ‐ ‐ For a perfect absorber 4.7 10 / on Earth’svicinity ‐ For a perfect reflector 2 ‐ Other perturbations ‐ Moon and Sun gravity fields, tides, Earth’s magnetic field T orbital period M satellite’s mass atmospheric density R orbit radius S projected satellite’s surface drag coefficient (typical 2.5) ∆ 3 ρ
  13. 13. 10 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 19 Orbital perturbations (ii): ‐ Accelerations caused by the main perturbation sources Comparsions of the disturbing accelerations for the main sources of perturbation Forteskue et al.,  T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 20 Orbital Perturbations (iii): ‐ Effects ‐ Oscillation of the orbital plane vs. nominal inclination ( Δ ) Δ e 2 J J cos ‐ Regression of the nodal line ∆Ω 3 1 1 ‐ Advance of the Perigee ∆ 3 2 1 1 4 5 sin Inclination= 63.4 → zero precession (Molniya orbit)
  14. 14. 11 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 21 LEO orbit: ‐ Polar LEO orbit: Global coverage (passes by the poles), typical altitude between 600 and 800 km ‐ Earth‐synchronous orbit: the ground track repeats its trace over the Earth at regular intervals The longitude shift ∆Ω = ∆Ω1 + ∆Ω2 in one orbit for the Ecuator is due to: ‐ The Earth’s rotation (dominant term) ∆Ω 2 [rad/orbit] Regression of the ascending node ∆Ω 3 [rad/orbit] T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 22 Earth’s rotation period: Te = 86164,09055+0,015 ts [s] ts: centuries from 1900 Te ~ 23h 56’ in relation to stars (24 h with respect to the Sun) An Earth‐synchronous orbit satisfies: n |ΔΩ| = m ∙ 2π , where n is the number of orbits m is the number of Earth revolutions (days)
  15. 15. 12 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 23 Sun‐synchronous orbit: the orbital plane revolves at the same angular rate as the Earth revolves about the Sun: 1 [rev/year]  ~ 1 [/day] ∆Ω 2 [rad/orbit] = 2π [rad/year]  Tes = 3,155815 ∙ 107 s (Earth‐Sun orbital period) The satellite flies over the territory at the same local time: ‐ 1 ascending pass: the satellite flies from South to North ‐ 1 descending pass: the satellite flies from North to South T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 24 Earth‐Sun synchronous orbit: Combination of the two previous requirements n |ΔΩ| = m∙2π n 2 2 2 → 1 1 ∆Ω 2 [rad/orbit] The orbit is indicated with indices n:m ∆Ω 2 (towards the West) Example: For a LEO: h ~550 – 950Km, T ~ 95‐100 min → ΔΩ~ 0,43 rad/orb → 2800 km separation between ground‐tracks at the equator
  16. 16. 13 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 25 Zero Drift orbit:  ‐ if m=1, there are n orbits that repeat every day: N:m hideal (km) 14:1 894 15:1 567 16:1 275 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 26 ‐ The separation between ground‐tracks is: ; ~6378 ‐ Usually, this separation is excessive for Earth observation: 14:1 → d=2862,43 km Example: LANDSAT 1,2 T=103,3 min 251:18 orbit → 251 orbits in 18 days (almost a 14:1 orbit: 251=14∙18‐1) Ψ=99 separation between tracks about 160 km 920
  17. 17. 14 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 27 ‐ Ground‐track examples 43:3 501:35 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 28 ‐ Visibility from the tracking ground station on the Earth: Limited visibility + limited bandwidth for downlink limit maximum amount of data that can be downloaded
  18. 18. 15 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 29 ‐ Artemis data relay satellite: http://spaceflightnow.com/news/n0303/20artemis/envisatartemis.jpg T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 30 • Geostationary orbit: Coverage with 1 satellite Global coverage with 3  geostationary satellites
  19. 19. 16 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 31 • Geostationary orbit: T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 32 • Geostationary orbit: ‐ If orbit inclination is not exactly 0,  inclination > 0 ‐> regression of nodal line (longitude decreases) inclination < 0 ‐> progression of nodal line (longitude increases) ‐ This results in an “8”‐like ground track (red line below) 1 arctan cos( ) 4      2
  20. 20. 17 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 33 • HEO orbits: Molniya orbit parameters: a = 26560 Km (T = 12 h) e = 0.722 (hp = 1000 Km, ha=39360Km) =270 (Southern hemisphere perigee)  arbitrary (depends on required coverage) T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 34 Other orbits: Solution of Newton’s equations 3 body problem: 2 rotating bodies around a third massive one m M2 M1 r1 r2 r baricenter L1 L4 L5 L2L3 Lagrangian or Libration points: equilibrium points L1‐3 unstable; L4‐5 stable Map the sky (L2), monitor the Sun (L1) …
  21. 21. 18 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 35 Reentry • Reentry depends basically on:  atmospheric density: as a function of orbital height, solar activity…  The larger the density, the larger the friction, the sooner the reentry  ballistic coefficient:  where M: mass, A: section, and Cd: drag coefficient.  The larger the ballistic coefficient, the later the reentry • Recomendation: satellites reentry < 25 years • Reentry calculations:  NASA DAS (Debris Assessment Software):  https://www.orbitaldebris.jsc.nasa.gov/mitigation/das.html  ESA DRAMA (Debris Risk Assessment and Mitigation Analysis) and  MASTER (Meteoroid and Space Debris Terrestrial Environment Reference)  https://sdup.esoc.esa.int/web/csdtf/home T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 36 Sample reentries computed with DAS rev. 2.02 for a 6U cubesat h = 450 km reentry < 2 years h = 550 km reentry ~ 7 years
  22. 22. 19 T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 37 • Appendix: Defining the orbit – NORAD TLE parameters NORAD Two‐Line Element Set Format Data for each satellite consists of three lines in the following format: AAAAAAAAAAAAAAAAAAAAAAAA  1 NNNNNU NNNNNAAA NNNNN.NNNNNNNN +.NNNNNNNN +NNNNN‐N +NNNNN‐N N NNNNN  2 NNNNN NNN.NNNN NNN.NNNN NNNNNNN NNN.NNNN NNN.NNNN NN.NNNNNNNNNNNNNN  Line 0 is a twenty‐four character name (to be consistent with the name length in the NORAD SATCAT). Lines 1 and 2 are the standard Two‐Line Orbital Element Set Format identical to that used by NORAD and NASA. The  format description is: Line 1 Column Description 01 Line Number of Element Data 03‐07 Satellite Number 08 Classification (U=Unclassified) 10‐11 International Designator (Last two digits of launch year) 12‐14 International Designator (Launch number of the year) 15‐17 International Designator (Piece of the launch) 19‐20 Epoch Year (Last two digits of year) 21‐32 Epoch (Day of the year and fractional portion of the day) 34‐43 First Time Derivative of the Mean Motion 45‐52 Second Time Derivative of Mean Motion (decimal point assumed) 54‐61 BSTAR drag term (decimal point assumed) 63 Ephemeris type 65‐68 Element number 69 Checksum (Modulo 10) (Letters, blanks, periods, plus signs = 0; minus signs = 1) Line 2 Column Description 01 Line Number of Element Data 03‐07 Satellite Number 09‐16 Inclination [Degrees] 18‐25 Right Ascension of the Ascending Node [Degrees] 27‐33 Eccentricity (decimal point assumed) 35‐42 Argument of Perigee [Degrees] 44‐51 Mean Anomaly [Degrees] 53‐63 Mean Motion [Revs per day] 64‐68 Revolution number at epoch [Revs] 69 Checksum (Modulo 10) All other columns are blank or fixed. Example: NOAA 14 1 23455U 94089A 97320.90946019 .00000140 00000‐0 10191‐3 0 2621  2 23455 99.0090 272.6745 0008546 223.1686 136.8816 14.11711747148495  https://www.celestrak.com/NORAD/elements/ T1. Introduction and Orbits IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 38 • Bibliography  Spacecraft System Engineering, P.  Fortescue, J. Stark, and G.  Swinnerd, 2003 http://www.satflare.com/track.asp#TOP
  23. 23. 1 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 11 IEEE Young Professionals in Space T2. Space Environment and Thermal Control T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 2 • Objectives of the lesson – Understand the Radiation in the Space Environment and how  it affects the Satellite Subsystems – Understanding heat modeling • Black‐body and Sun radiation – How to design satellites that • Avoid extreme temperatures • Avoid large temperature changes • Allow heat exchange
  24. 24. 2 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 33 Space Environment • Sun dominates the space environment of the Solar System • It is a thermonuclear fusion reactor • Surface is at ~5900 K • Solar atmosphere consists of  ‐ lower region extending to some thousand kilometers Temperature peaks at 10000 K Emits mainly UV rays ‐ upper region extending to several solar radii Temperature is around 2∙106 K,  but very variable X‐ray emission T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 44 Space Environment Its emission spectrum approximates to a 5900 K black body
  25. 25. 3 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 55 Solar Radiation: Sunspots and Variability • Sunspots are regions which are cooler than the surrounding surface. • The larger the number of sunspots, the larger the emission of  radiation associated with solar flares: radio, X‐rays, γ‐rays • ~ 11 year cycle T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 66 Solar Radiation: Effects on Materials • Polymers are particularly sensitive to high energy photons • Modifications in resistivity • Optical changes • Solar arrays are particularly sensitive to UV, coverglass and adhesive subject to darkening • May enhance erosion by atomic oxygen etc.
  26. 26. 4 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 77 Solar Wind • Solar Wind dominates space weather • Consists of a flux of charged particles at 200‐800 km/s • Earth is protected by its own magnetic field that deflects it  solar wind shapes it. • Particles trapped in the Earth's magnetic field “go” to the Van Allen  belts. • Only observable on Earth when it is strong enough to produce boreal  aurorae, geomagnetic storms or plasma tails in comets. T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 88 Van Allen Radiation Belts • Contain energetic particles trapped in the Earth’s magnetic field • Inner belt (100‐10000 km) has a larger concentration of protons • Outter belt (13000‐60000 km) consists mainly of electrons • The effects caused by these particles are degradation of electronics due to accumulated radiation dose degradation of solar array performance single event upsets (SEUs) dielectric discharging
  27. 27. 5 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 99 Relative radiation level experienced in different orbits (from STMicroelectronics)  Calculation can be done using software tolos such as:   SPENVIS (SPace ENVironment Information System) https://www.spenvis.oma.be/ T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 1010 South Atlantic Anomaly: the “Bermuda triangle of the spacecrafts” • Region where the inner Van Allen belt is closer to the Earth • Exposes satellites flying below the belt to higher radiation doses,  causing disruptions on systems. • Caused by the non‐concentrity of Earth’s magnetic field.
  28. 28. 6 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 1111 Geomagnetic Storms Disturbance of Earth’s magnetosphere as a consequence of solar ions caused by disruption Effects: Satellite systems: Particle damage, UV, Atmospheric drag Power Systems: Geomagnetically induced currents Navigation Systems: Changes in ionosphere Communication: Ionospheric irregularities T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 1212 Single Event Effect A Single Event Effect is caused by ions or electro‐magnetic radiation  striking a sensitive micro‐electronic device. • Single event upset (SEU): A change of state or transient • Sigle event Latchup (SEL): Event causes loss of device funcionality • Sigle event Burnout (SEB): Device destruction is caused due to a hight current state in a power transistor.
  29. 29. 7 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 1313 Examination of nearly 1300 single‐event upsets from one  computer on the TAOS mission shows that nearly 50 percent  occurred in the South Atlantic Anomaly, whereas only 5 percent  of orbital time was spent there.  T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 14 Thermal Control - Extreme Environment - Space ~ 3 K - Intense solar radiation - Eclipses - Vacuum: no air circulation ⇒ risk of overheating - Narrow temperature ranges for devices - 0 to 50C for batteries - ‐10 to 150C for most electronics - 7 to 35C for fuel - Large and fast thermal gradients - Alignment distortion - Fast expansion and contraction - Risk of material cracking
  30. 30. 8 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 15 Radiation – Sun emits approximately as a black body – A black body is an ideal body that… absorbs all the energy at all wavelengths from all directions and  polarizations, and in thermal equilibrium re‐emits it all following  Plank’s law ‐ Radiation equations 1. Wien’s law: energy peak at σ= Wien’s displacement constant (~2898 μm∙K) 2. Planck’s law: energy density by wavelength , 2 1 1 3. Stephan‐Boltzmann’s law (will be seen later) [http://www.oswego.edu/~kanbur/a10 0/lecture13.html] T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 16 • Radiation – Sun temperature: ~ 6000K Earth temperature: ~ 300K ⇒ λmax= 483 nm (visible, blue)  ⇒ λmax= 10 µm (thermal infrared) [http://scienceofdoom.files.wordpress.com/2009/11/blackbody_curve‐sun‐earth.jpg]
  31. 31. 9 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 17 • Radiation – A black body is ideal, the Sun is not – Real bodies have • Absortivity: α – fraction of energy absorbed, relative to that of a black body • Emissivity: ε – fraction of energy emitted, relative to that of a black body – Atmosphere also modifies spectrum Black‐body (blue) / Sun TOA (red) T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 18 • Radiation reaching the satellite [Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003] Typical Spacecraft Thermal Environment
  32. 32. 10 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 19 • Radiation reaching the satellite – Direct light from the Sun ~1400 W/m2 – Albedo (reflected light from the Sun on the Earth’s surface) ~30%  of direct solar radiation – ~10‐40% over soil – ~5% over water – ~40‐80% over clouds – Earth’s radiation ~250 W/m2 • Internal heat dissipation • Radiation emitted by the satellite – Because of body emission Etotal=σ∙T4 [W/m2] (Stephan Boltzmann’s Law) [https://engineerjau.wordpress.com/2013/06/09/where-does-the-stefan-boltzmann-constant-come-part-1-of-2/] T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 20 • Heat balance – Equilibrium: Qin = Qout – Body illuminated by the Sun: Pdirect + Palbedo + Peartshine + Pdissipated = Pemitted • For GEO orbit only direct light is significant – Body in shadow: Peartshine + Pdissipated = Pemitted – Depends on absorptivity and emissivity
  33. 33. 11 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 21 • Heat balance – Absorbed Pdirect =Direct solar∙Silluminated∙αs [W/m2] Palbedo=Albedo ∙Silluminated∙αs [W/m2] Pearthshine=Earthshine ∙Silluminated∙ε [W/m2] Pdissipated=Power  [W] – Emitted Pemitted = Stotal∙ε∙T4∙σ • Absorptivity and emissivity examples – White paint:  α=0.15,   ε=0.9 – Black paint:  α=0.9,   ε=0.85 – Aluminum:  α=0.15,   ε=0.05 – Gold:  α=0.25,   ε=0.04 – Solar cells (Si):  α=0.75,   ε=0.82 – Solar cells (AsGa): α=0.88,   ε=0.80 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 22 • Heat balance dynamics C: heat capacity  Transient‐state Oscillation in steady‐state
  34. 34. 12 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 23 • Example 1: – Orbit time 94 min, 63 min in light – Considering 1 kg of aluminum, 10 cm each side – 1 day simulation – Materials: Silver (α=0.37,  ε=0.44) ⟶ blue AsGa (α=0.88,  ε=0.80) ⟶ red Si  (α=0.75,  ε=0.82) ⟶ green Silver (blue, ‐7 to 3 °C), AsGa (red, ‐12 to 10 °C), Si (green, ‐17 to 3 °C) [°C] Time [∙104 s] T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 24 • Example 2 – Orbit time 94 min, 63 min in light – Considering 1 kg of aluminum, 10 cm each side – 1 day simulation – Materials: Gold (α=0.25,  ε=0.04)  Gold: 167 to 175 °C [°C] Time [∙104 s]
  35. 35. 13 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 25 • Thermal design (1/3) – Thermal control requires a balance of onboard heat emission and absorption – Radiation surfaces have to be sized large enough to keep temperature within upper temperature limit during solar exposition – Whole system has to be designed respecting security margins – Coated and painted surfaces have to be designed according to valid temperature ranges – Inner sides are black painted to maintain homogeneous temperature – Heaters have to be sized large enough to keep temperature above the lower temperature limit during umbra periods – Non‐radiating surfaces are generally insulated with multilayer insulation – It is highly recommended to simulate internal temperature gradient (Thermal Desktop, ESATAN…) T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 26 • Thermal Systems – Passive systems • No power consumption • High reliability • Low heat transfer capabilities (except heat pipes) – Active systems • Power and telemetry requirements • Lifetime considerations, operational risk • Adaptable, flexible and higher heat transfer capabilities
  36. 36. 14 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 27 • External Systems – Passive systems • Radiator surfaces • Coating (paint) [http://novitastechnology.com/Projects.html] T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 28 • External Systems – Passive systems • Thermal Blankets MLI: Multi‐Layer  Insulation [http://en.wikipedia.org/wiki/Multi‐layer_insulation]
  37. 37. 15 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 29 • External Systems – Passive systems • Sun Shields http://www.jpl.nasa.gov/images/aquarius/aquarius‐20090429‐browse.jpg T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 30 • External Systems – Active systems • Louvers
  38. 38. 16 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 31 • Internal Systems – Passive systems • Heat Pipes (high transfer capability) • Radiating systems [http://hectorleonosorio.blogspot.com/2010/07/tubo‐ de‐calor‐heat‐pipe.html] http://en.wikipedia.org/wiki/File:Laptop_Heat_Pipe.JPG http://www.cheresources.com/htpipes.shtml T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 32 • Internal Systems – Active systems • Heaters • Peltier cells [http://mechanical‐ engineering.esa.int/thermal/images/heater3big.jpg] [http://en.wikipedia.org/wiki/File:Peltier_(detail)_LMB.png]
  39. 39. 17 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 33 • Design Verification – Similarity • Use of space qualified materials – Verification Analysis • Comprehensive Thermal Math. Model (CTTM) – Thermal Vacuum Test • Temperature ranges • Temperature gradients • Validation of CTTM T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 34 • Example: SMOS at Large Space Simulator
  40. 40. 18 T2. Space Environment and Thermal Control IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 35 • Bibliography  o Spacecraft System Engineering, P.  Fortescue, J. Stark, and G.  Swinnerd, 2003 o ECSS‐E‐ST‐10‐04C Space Environment (15 November 2008) in Space Engineering  o ECSS‐E‐ST‐10‐12C Method for the calculation of radiation received and its  effects, and a policy for design margins (15 November 2008)  in Space Engineering  o ECSS‐E‐ST‐20‐06C Spacecraft charging (31 July 2008) in Space Engineering 
  41. 41. 6/26/2018 1 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 1 IEEE Young Professionals in Space T3. On Board Data Handling T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 2 Objectives of the lesson: ‐ Understand what “On Board Data Handling” is ‐ Understand what kind of telemetry data is required ‐ Architecture of the OBDH ‐ Issues affecting the OBDH design
  42. 42. 6/26/2018 2 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 3 Introduction T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 4 • Involved subsystems. • Types of data, requests, actions…. • Flight software and On‐Board Computer: architecture, OS, CPU. Payload Recording device EPS, ADCS… Communications Subsystem On-Board Computer TT&C Ground segment Payload data Configuration Logs,configuration Firmwareupdates Mission control, Sw. updates Digital buses Introduction (cont’d)
  43. 43. 6/26/2018 3 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 5 Telemetry data (downloaded to GS): • Housekeeping data:  Monitor the status of the s/c.  Sensors, device states, system  logs (e.g. error messages). • Attitude data:  Determine/monitor attitude of  s/c.  Mode of operation, controller  state, sensor readings. • Payload data:  Payload data is needed to  monitor the status of the  payload and the experiments to  be processed before being  transmitted to the ground. Spacecraft operations Telecommands (uploaded to s/c): • Control modes of operation:  Perform initial commissioning.   Enable states in spacecraft FSM.  Handle contingency situations.  Change payload modes or setup. • Configure devices and subsystems:  Change parameter values (they  have been designed to allow that).  • Fix errors (to some extent):  Upgrade software and firmware.  Access digital buses and  microcontrollers. T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 6 Telemetry data – housekeeping data • Data volume is critical (constrained bandwidth, data‐rate and  visibility). • Number of different parameters to monitor. • Low sampling frequency (period from 1 s to 2min) • Telemetry data can be transmitted to ground • in real time (e.g. beacon signals) • stored (delayed transmission in case of link unavailability)  THE 3CAT‐1 CASE: 9600 bps, MTU: ~64 kB, Selective repeat ARQ protocol  → 1 min (best case, no errors in protocol/handshakes.)
  44. 44. 6/26/2018 4 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 7 Telemetry data – housekeeping data • Data volume is critical (constrained bandwidth, data‐rate and  visibility). • Number of different parameters to monitor. • Voltages, Current Consumption, ADCS… • Low sampling frequency (period from 1 s to 30 s) • Telemetry data can be transmitted to ground • in real time (e.g. beacon signals): • Very useful for amateur bands, • stored (delayed transmission in case of link unavailability): • Just for critical information that cannot be lost. THE 3CAT‐2 CASE: 9600 bps, MTU: ~26 kB file chunks, Selective Repeat ARQ file oriented  protocol  45sec/chunk (medium case, assuming 5% of PER, packet  size=125 bytes) T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 8 Telemetry data – housekeeping (example) Type Source Analog / Digital Temperature Solar panels, batteries,  OBC Analog (thermistors,  RTD, lcs,…) Voltages & currents Power supplies solar  panels, power buses Analog (sensing  circuitry) Operation modes Payload status, heaters  (on/off) Digital Actuation / deployment  mechanisms Solar panels and  antennas deployment  Digital
  45. 45. 6/26/2018 5 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 9 Telemetry data – housekeeping data, considerations • Number of required acquisition channels • Accuracy required • Precision of any telemetry channel shall allow an assessment of  the nominal or out‐of‐tolerance status of the monitored  parameters • Sufficient telemetry shall be available to allow the verification of: • Power budget with accuracy better than  e.g. ±5% • Temperature with an accuracy better than e.g. ±1% • Types of sources. • Format and storage (standardization is critical):  • e.g. Analog voltage 0 to 5 VDC: units are [V] or [mV]? • Storage type: 8/16/32‐bit integer, floating point, ASCII (“5123” =  5.123 V) THE 3CAT‐1 CASE: EPS HOUSEKEEPING Data volume without overheads ~63.3 kB = ~8000 registers (timestamp + data).  5 sensing magnitudes (up to 7 channels) @ Ts = 1 min. → < 5 hours of history. T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 10 Telemetry data – housekeeping data, considerations • Number of required acquisition channels THE 3CAT‐2 CASE: EPS HOUSEKEEPING Three different logs ~96 kB @ ~38 bytes/register = ~2500 registers Compressed  Down to 20kB file size (high compression ratio) Lots sensing magnitudes at a rate of 20 seconds  ~12‐13 hours  1 log file !!!!  Which is the revisit time over the Ground Station !!!! THE 3CAT‐2 CASE: ADCS HOUSEKEEPING Three different logs ~96 kB @ ~84 bytes/register = ~1150 registers Compressed  Down to 40kB file size (medium compression ratio) Lots sensing magnitudes at a rate of 5 seconds  ~1.5 hours  1 log file !!!! Which is the orbital period !!!!
  46. 46. 6/26/2018 6 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 11 Telemetry data – attitude determination • Attitude determination can be based on: ‐ Sun and Earth sensors ‐ Accelerometers ‐ Gyros ‐ Magnetometers ‐ Star trackers • Few data is normally required • Sampling rate Normally low sampling rate (T = 1~ 60s) even if it can be higher on  demand  (ex.: stabilization of the spinning) T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 12 Telemetry data – payload data • Data volume is variable. Different payloads will generate different  amounts of data. • Ultimately limited by the link budget (not the storage). • Sampling rate: • May be constant and regular on time (ex.: to study a degradation  in space conditions) • Fast sampling bursts (ex.: to study fast dynamic changes of a  MEMS experiment) • Experiments may be disabled after their duty cycle in order to  release storage regions to other tests. • Payload data is not necessarily kept forever in the recording device.  Downloading it may mark the recording device region as “available”  again.
  47. 47. 6/26/2018 7 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 13 Telecommand data (example) DESTINATION EXAMPLES Spacecraft Orientation (activation of wheels & gyros) Micro‐propulsion systems Manual activation of mechanisms (heaters) C&DH unit Changes of operative modes (stand‐by, energy saving,  active modes) Data upload/download Power section Change on battery level based operation modes Reset RF Rx/Tx Beacon On/Off Payloads Power on/off, sampling rates, activation timer, … T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 14 Telemetry data packet • Telemetry data is encapsulated into a packet and prefaced by a packet  header • Packet header contains: ‐ Source ID ‐ Number of the current packet in the sequence of packets produced ‐ Length of the data field ‐ Packet error control can be added.
  48. 48. 6/26/2018 8 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 15 Telecommand data – packet telecommand layers • Packetization layer ‐ TC packet consists of a fixed length packet header and a variable  length data field ‐ Header contains: destination ID, packet length, packet control  error features (optional) • Segmentation layer Combination of multiple TC packets with data contained inside • Coding layer  Provides the error correction and synchronization required by the  spacecraft • Physical layer RF link itself T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 16 Telecommand data • Elementary commands or complex sequences (SW update) • Validated by the OBDH unit before its execution ‐ OBDH transfers the commands, normally to the OBC ‐ Internal communication protocols (I2C, SPI, RS232, …) ‐ OBDH confirmation to the ground of the data received before its  execution. • Telecommand can be executed instantly or deferred
  49. 49. 6/26/2018 9 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 17 Spacecraft Flight Software: general architecture. • Common topologies: centralized or distributed. C&DH Subsys. 3 Subsys. 1 Subsys. 4 Subsys. 2 Subsys. 4 Integral modules (hardware)  connected through digital buses or  software modules interfacing with  OBDH core. Subsys. 1 OS Framew ork Subsys. 1 OS Framew ork Subsys. 1 OS Framew ork Spacecraft Bus e.g. Voyager, Galileo (NASA’s outer  planetary spacecraft), AAUSAT3  (nano‐satellite mission) T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 18 Spacecraft Flight Software: • Structured software that is: • Reliable:  • Precludes error propagation from one module to others; • Prevents deadlocks; • Controls timeouts, watchdogs, event sequencing, transitions to safe  states; • Detects invalid input data; • Protected memory regions (isolation); • … • Automatically analyzable: verify code flow, data integrity, deadlocks,  infinite loop conditions, etc. before compile‐time! (Software verification  processes are usually hard and tedious.) • Traceable: it should be easy for ground operators to determine what is  happening with reduced information.  • Reusable: standard interfaces, modular architectures, encapsulation of  components. • Modifiable (at design‐time), and updateable (at run‐time!) Specially valuable when missions have short development cycles (e.g. nano‐satellites)
  50. 50. 6/26/2018 10 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 19 Spacecraft Flight Software: • Implementation of multiple‐level Finite State Machines. • High‐level mission states → internal sub‐states → subsystem modes  → device commands. • Should always guarantee spacecraft safety, i.e. tend to be complex  and should be automatically analyzable.  • E.g. if contingency:  SYSTEM RESET (may take from several seconds to some hours), SAFE STATE (only critical subsystems are enabled). • E.g. 3Cat‐1 energy and system FSM (each high‐level state is  decomposed into several routines): CONTINGE NCY STANDBY PAYLOADS COMMS IDLE INIT LE ∨ IE LE ∨ S LE ∨ S Controlled by EPS OBC shut down Energy critically low OBC shut down Low-power mode Comms RX enabled Payloads disabled Nominal power mode Comms. full mode (RX & TX) Payloads enabled 0% 10% 15% 75% 100% OBC power-off OBC power-on State of charge T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 20 Spacecraft Flight Software: • On‐board pre‐processing of data: only valuable data is downloaded. • Failure Detection, Isolation and Recovery techniques, at software  level. Systems that are capable of identifying or recovering from  errors without orders from ground operators: • E.g. when round‐trip delay times are too long to wait for the  ground operators to react. • Current trends: autonomous spacecraft.  • “Minimum” human intervention: on‐board Mission Planning and  Scheduling systems. • On‐board identification of interesting data acquisition  opportunities.
  51. 51. 6/26/2018 11 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 21 Spacecraft Flight Software: • Real‐time operating systems (RTOS): • Deterministic latencies RT scheduler:  predictable execution. • RT ≠ fast compu ng. • Priority‐based. Preemptive / non‐preemptive. • Programming aspects: kernel manages Real‐ Time Tasks (similar to threads). • Examples of RTOSes used in spacecraft: (*) http://opensource.gsfc.nasa.gov/projects/cfe/ • Common frameworks and standards: • NASA cFS/cFE (Open Source, Free License) (*). Available for RTEMS, Linux, VxWorks… • TSP: Time and Space Partitioning (Avionics Application Standard Software Interface  ARINC 635 standard.) Task  (NAME / PRIORITY) A / 1 B / 2 C / 3 Execution on CPU Task context switches are managed  by the kernel scheduler (i.e. they  are deterministic, predictable).  VxWorks: proprietary.  RTEMS: open source, free license.   Micrium uC/OS‐III: open‐source.  QNX: proprietary.  LynxOS: proprietary.  FreeRTOS: open‐source, free license.  Xenomai (Linux patch): soft‐real‐time,  open‐source, free license.  PREEMPT_RT (Linux patch): soft‐real‐ time, open‐source, free lic. T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 22 CPU and OBC boards Commercial-of-the-shelf components are used (for CubeSats). Pumkin FM430 http://www.cubesatkit.com Nanomind A712C http://gomspace.com/ Intrepid http://tyvak.com/products/products.html CPU MSP430, 16bit @8 MHz ARM7, @ 8-40 MHz ARM9, 32bit @ 400 MHz Memory 55 kB Flash, 5 kB RAM 2MB RAM, 4 MB Code storage, 4 MB data flash 64 MB SDRAM, 128 MB NAND Flash, 128 byte EEPROM CPU Power consumption From 0.2 μA to 330 μA (@1 MHz) 3.3 V, 70 mA (nominal) From 0.9 μA to 80 mA OS SalvoOS, FreeRTOS FreeRTOS OS Linux 2.6.x, OpenEmbedded Integrated peripherals Direct Memory Access controller, 12bit ADC and DAC, 2 Timers with pulse with modulation, 2USART with SPI, I2C or Asynchronous UART, 6 8bits I/O ports CAN and I2C interfaces RTC - real time clock On-board magnetometer, 3x PWM drivers, 6x sun-sensor inputs, 3x Rate-gyro inputs… 10/100 Ethernet, USB2.0, 2 USB Host, 1 USB Device, 1 Synchronous Serial Controller, 1 SPI, I2C, 64 digital I/O ports, 16bit parallel bus, 5 serial interfaces (UART/USART), External SD Card interface, 4 10-bit ADC
  52. 52. 6/26/2018 12 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 23 Issues affecting OBDH (i) • Radiation: • Total dose: for a unit installed externally of the ISS a total dose of 3 kRad for a 3‐years mission is foreseen with a box thickness of 3 mm Al  equivalent, safety margins applied • Total dose mainly due to X‐rays and Gamma‐rays.  • Components sensitive to total dose present functionality anomalies  (power consumption abnormal increase, ….) • Single event effects:  • SEE: generalized category of anomalies resulting from a single ionizing  particle  • SEU: Single event upset (non‐destructive ), SEL single event latchup (destructive), SEGR single event gate rupture (destructive), … • Radiation effects on design: • Radiation limits the number and the type of components available to the  designer : uController, CPUs, FPGA Field Programmable Gate Arrays ,  EEPROMs, Amplifiers, Voltage regulators, etc T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 24 Issues affecting OBDH (ii) • Radiation: solutions • Radiation‐hardened devices: • Tolerate higher radiation doses. • Special manufacturing techniques.  • Bipolar technology (not CMOS) • Shield components. • On CPU: usually SPARC or PowerPC architectures. • E.g. SPARC‐V8 LEON3 (VHDL‐synthezisable, free license),  project started by ESA. • Error‐Correcting Codes, e.g. NAND‐Flash with 4‐bit BCH ECC. • Redundancy: at device‐ or circuit‐level. (Software redundancy could  also be possible) Block #1 Block #2 Block #3 Input data Arbitrator (voting) Input data 1011 1001 1011 1011 1011
  53. 53. 6/26/2018 13 T3. On Board Data Handling IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 25 Issues affecting OBDH (iii) • Thermal Constraints: • Thermal environment and vacuum limits the types and number  of components available to the designer.  • Typical operational temperature range for a spacecraft OBDH  unit could be  –20 °C to +50 °C • Electronic components are usually classified and sold in three  different families: • Commercial range: 0 to +70 °C ( plastic package) • Industrial range : –40 to +80 °C  • Military range : –55 to +125 °C ( ceramic package) [sometimes for automotive applications as well] • Heat from electronic components can be removed only by  conductive paths (heat pipes, no fans, no finned heat sinks....)  or by radiation
  54. 54. 6/26/2018 1 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 1 IEEE Young Professionals in Space T4. Attitude Determination and Control System T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 2 Objetives of the lesson: ‐ Payload and sensors requirements for operation (e.g. Telescope) ‐ Mission operation pointing requirements ‐ Power demand: Sun‐pointing ‐ Communications: antennas pointing ‐ Thermal regulation: avoid heat gradients, heat dissipations to deep space ‐ Thruster pointing for orbital change ‐ Mission stat
  55. 55. 6/26/2018 2 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 3 Basics: Linear Momentum • Center of mass M∙roc = Σ m∙rOP Lets us to handle withe the full body as an equivalent particle (e.p)  of mass M in the position roc. In that position Σ(m∙rcp) = 0 is  acomplished always. • Linear momentum L = M∙v • Basis to translation/orbit dynamics • Impulse L 0 • An impulse is equal to the change in the momentum that it changes T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 4 Angular Momentum Moments of inertia: Ixx, Iyy, Izz e.g. Ixx =  (y² + z²)∙dm  Measures the distance from an axis: (y² + z²) = d² from x‐axis Products of inertia: Ipq e.g. Ixy =  xy∙dm  Matrix of inertia As an equivalent to the total mass of the body, it relates the  angular momentum with the angular speed vector
  56. 56. 6/26/2018 3 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 5 Angular Momentum Torque Impulse T 0 Equivalent to the linear momentum, but since it is not just  proportional to the total mass, but it depends on a matrix with products of inertia acting as cross‐coupling among diferent axes, it is not as easy understandable.  The form: ω] = [Ic]∙T , helps to understand it. T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 6 Momentum Build‐up • As in the linear case, only external torques affect the total angular  moment of the system because Σ Text = 0 → H is constant • External torques could be either on board torquers, such as thrusters or magnetorquer rods, or external disturbances (eg. Solar wind). • Whether the spacecraft has a momentum or not, Attitude Control  System (ACS) has to manage the momentum in order to get the desired performance or to damp the external disturbances
  57. 57. 6/26/2018 4 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 7 Attitude motion of specific types of spacecraft Spacecraft Without momentum bias (MB)       With momentum bias 3 axis stabilized (no MB) Spinner Hybrid Partially de‐spun 3 axis stabilized (with MB) T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 8 Momentum bias As in the linear case with L, momentum reduces ther sensitivity to torque → gyroscopic rigidity. ψ Momentum bias is used by spinning‐up the spacecraft before rocket firing to avoid misallign with course. Then, it is usually spun down. When a torque is applied to a biased body, it introduces a nutation mode. ACS must damp it. When the torque stops, it introduces more nutation oscillation, a good timing has to be set in order to compensate the new one with the former.
  58. 58. 6/26/2018 5 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 9 Spacecraft →  With momentum bias →  Spinner METEOSAT‐1 METEOSAT‐3 Font: http://es.wikipedia.org/wiki/Archivo:METEOSAT.gif MSG‐3 is the third in a planned series of four satellites operated by  Eumetsat, the European Organisation for the Exploitation of  Meteorological Satellites Read more at: https://phys.org/news/2012‐03‐msg‐satellite‐ready‐ weather‐monitoring.html#jCp T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 10 Simple spinner Payload and other subsystems are placed in the spinning section • Power: Solar array limited to sides • Thermal: avoids heat gradients • Communications: non directive anthennas • Limited mounting space for non‐scanning payloads • Spin Axis: Least intertia axis could only be chosen for a limited duration (rockets). For a satellite demanding a preservation of the spinning attitude, maximum inertia axis must be chosen. • Nutation: spinning introduces an oscillatory mode.   requires the use of passive dampers or control torquers
  59. 59. 6/26/2018 6 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 11 [https://www.youtube.com/watch?v=R8jiU1EYIAM] Spacecraft With momentum bias → Spinner → Partially de‐spun GIOTTO T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 12 Dual spinner Payload, antenna and feed are placed on a de‐spun platform • Power and thermal control similar to the simple spinner • Spinning parts inertial axes should be balanced • As a spinning body, nutation occurs and has to be controlled by ACS
  60. 60. 6/26/2018 7 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 13 Spacecraft With no momentum bias (no MB) → 3‐axis stabilized T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 14 Image font: http://www.esa.int Spacecraft With no momentum bias (no MB) → 3‐axis stabilized e.g. SMOS (and most other Earth Observation satellites)
  61. 61. 6/26/2018 8 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 15 Image Font: http://www.ask.com /wiki/Magellan_probe Dual spinner • Instead of spinning the spacecraft, wheels allow to  keep it stabilize it. • Wheels store momentum and changes satellite facing by varying its speed.  • 3‐axis wheel stabilization allows control almost independently for each axis by placing them ortogonally.  • Usually, a 4th wheel is added in a combination affecting all of the axis in order to have redundancy and reliability. • Momentum wheels accomodate fluctuations of 10% of the bias T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 16 ACS Block Diagram (i) ‐ Attitude and Control System keeps the spacecraft in a given attitude.  ‐ Compensates all possible disturbances appropriately. ‐ Furthermore, it has to let the spacecraft perform the attitude changes for  the correct developement of the mission. Typical example is to set the pointing to the Earth or the Sun. ‐ Desired torques combined with disturbances affect the spacecraft and are  sensed. ‐ On‐board and on‐ground computers compute the torque to be applied to  correct and/or change the attitude following the established plan. Image font: Spacecraft Systems Engineerig. Fortescue, Stark, Swinerd
  62. 62. 6/26/2018 9 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 17 ACS Block Diagram (ii) Control loop:  PID control is by far the most common Proportional:  proportional to error good response to quick disturbances Integrator:  proportional to the accumulated (“integrated”) error good to remove biases Derivative:  proportional to rate of change of error good for stability Pointing and momentum management are usually controlled independently T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 18 External Torques Disturbances • Aerodynamic < ~500 km • Magnetic ~ 500 – 35000 km • Gravity gradient ~ 500 – 35000 km • Solar radiation > ~ 600 km Controllables • Gas Jet • Magnetorquers • Adjustable geometry
  63. 63. 6/26/2018 10 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 19 Gas Jets • Requires fuel • Suitable for any torque size • ON‐OFF control only T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 20 Magnetorquers • No fuel, but power required • Complicated Earth field • No torque about field line https://www.cubesatshop.com/product/isis‐magnetorquer‐board/
  64. 64. 6/26/2018 11 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 21 Adjustable geometry • Require fuel • Suitable for any torque size • ON‐OFF control only T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 22 Internal Torques - Internal Disturbance Torques - Mechanisms (deploying parts) - Fuel movement - Controllable Internal Torquers - Reaction wheels - Momentum wheels
  65. 65. 6/26/2018 12 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 23 Attitude Sensors - In order to keep a good attitude, attitude must be measured. - Reference sensors, with external information sources - Stars (1 arcsec) - Sun (1 arcmin) - Earth (horizon) (6 arcmin) - Magnetometer (30 arcmin) - GPS (6 arcmin) - Inertial sensors T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 24 Attitude Sensors ‐ Reference sensors, with external information sources ‐ Stars (1 arcsec) ‐ Fixed‐head star tracker: electronically scanned imatge . CRT technology . High power and mass . Magnetic interference
  66. 66. 6/26/2018 13 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 25 Attitude Sensors . Reference sensors, with external information sources . Stars (1 arcsec) . Star scanner ‐ No moving parts ‐ Mounted on spinning spacecraft . CCD Star tracker (and mapper) https://www.cubesatshop.com/product/mai‐ss‐space‐sextant/ https://www.cubesatshop.com/product/nst‐1‐nano‐star‐tracker/ T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 26 Attitude Sensors . Reference sensors, with external information sources . Sun Sensors . Single axis (and two axis by orthogonal combination) . Two‐axis CCD array Sun sensor https://www.cubesatshop.com/ product/nss‐cubesat‐sun‐ sensor/ Photodiode (1 per face) 2‐Axis digital sun sensor https://www.cubesatshop.com/product /ssoc‐d60‐2‐axis‐digital‐sun‐sensor/ https://www.cubesatshop.com /product/mai‐ke‐sun‐sensor/ 1‐Axis digital sun sensor
  67. 67. 6/26/2018 14 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 27 Attitude Sensors . Reference sensors, with external information sources . Earth‐horizon . Infra‐red spectrum sensor . Target always present . Slower response https://www.cubesatshop.com/product/mai‐ses‐ir‐earth‐sensor/ T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 28 Typical Satellite Attitude States and Transitions (i) • After deployment the satellite (CubeSat) is completely disconnected for  TBD minutes in order to avoid any perturbation to  the main passenger. • Detumbling is responsible to slow‐down satellite oscillations and stabilize  it using 3‐axis magnetorquers: B‐dot algorithm. 
  68. 68. 6/26/2018 15 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 29 Typical Satellite Attitude States and Transitions (ii) • Once the derivative of the magnetic field intensity is smaller than a  predefined value ( ∗ ), it will transition to another mode if o Nominal if the EPS battery level is above a given threshold ( ∗ ), or  o Sun safe if the EPS battery level is below a given threshold ( ∗ ).  • In nominal mode reaction wheels are activated and will point e.g. the  antenna boresight or the camera optics to nadir or to the target. • However, if the above conditions are not met anymore, a return to the  “Detumbling” • Transition from Sun Safe Mode (Sun pointing) to Nominal Mode occurs  when the EPS battery level is above a given threshold ( ∗ ), and vice‐versa  when the EPS battery voltage is below a TBD % of its maximum value (TBC). • Transition from Sun Safe Mode to Survival Mode occurs when the EPS  battery level is below a TBD % of its maximum value (TBC). T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 30 All transitions are commanded from ground  (except transitions to emergency states) Typical Satellite Attitude States and Transitions (iii)
  69. 69. 6/26/2018 16 T4. Attitude Determination and Control System IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 3131 Bibliography Spacecraft System Engineering, P. Fortescue, J. Stark, and G. Swinnerd, 2003
  70. 70. 1 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 1 IEEE Young Professionals in Space T5. Electrical Power Supply System T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 2 • Motivation – All systems need energy to operate – Solar energy is unlimited and cheap, but not available for all missions – Sometimes a secondary power supply is need  • Learning objectives – Know different energy‐supply options Understand pros and cons of each option – Know how to dimension the energy system – Understand that performance degradation requires oversizing the  whole system – Know the typical blocks of an EPS
  71. 71. 2 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 3 • Design process – Power System depends on • Mission type (Earth orbit, inner planets, outer planets…) • Mission life • Mission payload – Power System Block Diagram Primary  Source Primary  Source Power  Management Batteries (Secondary  Source) Power  Regulation Regulated  Bus Primary  Sources T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 4 • Power Sources – Main • Solar Cells • Fuel Cells • Radio‐isotope Thermoelectric  Generator (RTG) • Nuclear fission systems • Solar Heat Systems – Secondary • Batteries Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003 System W/kg Power Available Cost $/W Solar Array 67 25 kW 600 R.T.G. 5 285 W 50‐60K Solar Dynamics 25 > 20 kW ‘High’ Fuell Cell 110 12 kW 40 Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003 Power Sources Performance Comparison
  72. 72. 3 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 5 • Solar panels – Group of Solar cells [http://photojournal.jpl.nasa.gov/catalog/PIA14172]v T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 6 • Solar Cell – Is a p‐n semiconductor device that converts photons in electrons • With no illumination, the p‐n junction achieves an equilibrated  state ⇒ no current flow • With illumination, photons with enough energy (gap band)  create electron‐hole pairs that create a potential difference in  the cell’s terminals ⇒ current flow http://physics‐tutor.site90.net/drupal/node/2solarcellcentral.com
  73. 73. 4 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 7 • Solar panel components – Gap band is the minimum energy amount needed to allow  photoelectric effect • It depends on material properties – Energy depends on frequency because of Plank’s law λ Where h is Plank’s constant (6.62∙10‐34  J∙s) • ⇒ Each material converts photons to energy until a maximum  wavelength Material Band Gap (eV) Maximum Wavelength (μm) Si 1.12 1.12 CdS 1.2 1.03 InP 1.344 0.923 GaAs 1.35 0.92 CdTe 2.1 0.59 GaP 2.24 0.554 Spacecraft System Engineering, Fortescue, P., Stark, J., Swinerd, G. pag. 329 http://www.ioffe.rssi.ru/SVA/NSM/Semicond/InP/bandstr.html T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 8 • Solar Cell Electrical Model Diagram IPH D RP RS ‐ ‐
  74. 74. 5 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 9 • Solar Cell Electrical Model – Current through load is maximum when ZLoad≈0 (shortcircuit) • Imax≈IPH – Voltage is maximum between terminals when ZLoad ≈∞ (open circuit) • Vmax ≈IPH∙RP – Transferred power is maximum when load is adapted to cell • Solar Cell P‐V & I‐V graph example: Spectrolab Blue: I(V[Volts]) [mA]  Red: P(V[Volts]) [mW] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 10 • Solar Cell P‐V & I‐V graph example – We can see clearly the Maximum Power Peak (MPP) – It depends on load impedance and incident power: • With less incident power, output voltage and current will be lower – It is recommended to set working point at MPP to get the maximum  power available • Use of Maximum Peak Power Tracking (MPPT) – Available power on solar cells depends on • Light power density (∼1400 W/m2) • Conversion efficiency – Manufacturing material – Temperature – Absorbed radiation • Solar cell surface
  75. 75. 6 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 11 • Solar Cell Material Comparison Material Si GaAs InP Efficiency 15,5% 19,5% 16,5% 10 years GEO EOL 8,8% 13,6% 16,2% Power loss/°C 0,438% 0,162% 0,206% Mass density 0,55 kg/m2 1,69 kg/m2 1,77 kg/m2 Relative Cost 1 5 36 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 12 • Radiation damage in solar cells – Radiation damages cells • Efficiency decreases • Less available power  [From Solar Cell Array Design Handbook by Rauschenbach, H. S.]
  76. 76. 7 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 13 • Temperature effect on solar cells – If temperature raises, efficiency decreases [From Solar Cell Array Design Handbook by Rauschenbach, H. S.] [http://www.altestore.com/howto/Solar‐Power‐Residential‐Mobile‐ PV/Off‐Grid‐Solar‐Systems/Electrical‐Characteristics‐of‐Solar‐Panels‐ PV‐Modules/a87/] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 14 • Solar array configurations: – One single solar cell does not give enough current and voltage  for almost any application – Solar cells can be connected in serial or parallel • Serial: output voltage is the sum of cell voltages • Parallel: output current is the sum of cell currents • In both situations, output power is the sum of powers of  the different cells
  77. 77. 8 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 15 • Solar array forming problems – In parallel blocks, if one cell is not  illuminated, its voltage will be 0 V  Current of other cells will flow through  that cell  Block voltage will be 0 V (short‐circuit)  Cell damage risk + ‐ Illuminated  solar cell Not  illuminated  solar cell ‐ In serial blocks, if one cell is not  illuminated, its current will be null  Current of other cells will not flow  trough it  Output current is null + ‐ Illuminated solar  cell Not illuminated  solar cell If a solar cell fails, it can act as: Short‐circuit: disables all cells set in parallel Open circuit: disables all cells set in serial Solution: add diodes T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 16 • Serial solution with diodes • If one cell fails and  becomes an open‐ circuit, current will  flow through the  diode. • If one cell becomes a  short‐circuit, diode  will not affect.‐ • Parallel solution with diodes • If one cell fails and becomes  an short‐circuit, other cells  will not be short‐circuited  because of this failed cell. • If one cell becomes a open‐ circuit, diode will not affect. ‐ … • This block has open‐circuit and short‐circuit protection • Can be set in parallel with other blocks (more current) ‐ If one of the blocks have lower voltage (e.g. no light), current will  not flow through it (without damage cell risk) • Can be set in serial with other blocks  (more voltage) ‐ If one of the blocks has lower current, it will flow through diodes  (without damage cell risk)
  78. 78. 9 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 17 • Solar Cells – Pros • Unlimited availability • Low failure risk • It can be built with common materials (Si) Cheap – Cons • Low efficiency (up to 19%) • Low power (oversized solar panels for high power consumption) • No energy while in eclipse (shadow )   Oversized solar panels  Large size ⟹ Low natural resonance frequency • Panel degradation (oversized solar panels) • Only for inner planets missions • Deployment mechanism complexity • Source of single point failure T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 18 • Low efficiency solutions (1/4)
  79. 79. 10 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 19 http://www.optoiq.com/index/photonics‐technologies‐applications/lfw‐display/lfw‐article‐display/314437/articles/laser‐ focus‐world/volume‐43/issue‐12/features/conference‐preview‐photonics‐west‐2008‐light‐it‐up‐and‐they‐will‐ comex2026‐to‐photonics‐west.html • Low efficiency solutions (2/4) – Multi‐junction Solar Cell • Each layer absorbs a wavelength band • Efficiency up to 42.3% T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 20 • Low efficiency solutions (3/4) – Back‐surface reflectors (BSR) to reflect unabsorbed radiation and  reduce heating • Used on Spot and Orion satellites – Using a p+‐region (BSF) at the rear of p‐regions to enhance  carrier  collector efficiency, but with high radiation fluency damage • Used on Hubble and Envisat satellites – Use of a textured front surface to reduce reflection from the cell  surface
  80. 80. 11 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 21 • Low efficiency solutions (4/4) Historic summary of champion solar‐cell efficiencies T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 22 • Fuel Cells – Designed for manned Apollo program – Use chemical energy converting a oxidation reaction into  electrical energy with minimal thermal changes • Example: Hydrogen‐oxygen fuel cell, which produces  water. Useful for manned missions.  H2O fuel cell schematic Direct Energy Conversion, Angrist, S. W., 1982
  81. 81. 12 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 23 • Fuel cell available energy – Available energy depends on ∆ – ∆ is “Gibbs (free) changing energy” or “free enthalpy” occurring in  the reaction,  – n is the number of electrons transferred, and  – F:  Faraday constant (9.65∙104 C/mol) • For H2O2, ∆ = ‐272,9 kJ/mol and n = 2. The resultant voltage is  1.229 V (considering an ideal cell) T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 24 • Typical hydrogen‐oxygen voltage‐current curve [Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003]
  82. 82. 13 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 25 • Fuel cell material comparison [http://www1.eere.energy.gov/hydrogenandfuelcells/fuelcells/pdfs/fc_comparison_chart.pdf] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 26 • Fuel cell pros – Clean – High efficiency – Cheap – Resultant substance can be useful • In hydrogen‐oxygen cell, the output is water • Fuel cell cons – Low power available – Short durability (less than a month) • Not suitable for outer planet missions
  83. 83. 14 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 27 • Radio‐isotope Thermoelectric Generator (RTG) – For outer planet missions • Cannot be used: – Solar panels: Farther than mars, panels have to be extremely  large – Fuel cells: Their durability is shorter than a month (in most  cases, not enough time to arrive) • Requirements: – High durability (Galileo took 6 years to arrive to Jupiter and  Voyager 2 took  12 years to arrive to Neptune) • Solution: nuclear energy T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 28 • Radio‐isotope Thermoelectric Generator (RTG) – Based on thermoelectric effect • It is possible to generate a voltage between 2 materials if there  is a temperature gradient between them – RTG systems use two semiconductors (p and n) with a junction to  exploit this effect • Low efficiency (<10%) ⇒ Need to remove wasted heat [Direct Energy Conversion, Angrist, S. W., 1982]
  84. 84. 15 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 29 • RTG available power – The output power depends on junction temperature • A radioactive material heats the junction while it decays • As it decays, the available power decreases . / where τ1/2 is half‐life – The lower τ1/2, the higher the power density Example: Cassini‐Saturn design life was 11 years. After this time, required  power was 628 W. The only possible fuel was Plutonium. Isotope Fuel Form Power Density (W/g) τ1/2 (years) Polonium 210 GdPo 82 0.38 Plutonium 238 PuO2 0.41 86.4 Curium 242 Cm2O3 98 0.4 Strontium 90 SrO 0.24 28.0 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 30 • RTG pros – Independent of orientation and shadowing – Independent of distance from the Sun – Low power levels for long time periods • Example: Voyager‐2, operative from August, 1977 – They are not susceptible to radiation damage • Example: Van Allen Belts, outer space missions… – Suitable for missions with long eclipse time • Example: lunar landers
  85. 85. 16 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 31 • RTG cons – It has to be deployed away from the satellite • Radiation affects spacecraft electronics • Need high temperature operation • Example: Galileo, RTGs deployed on lengthy booms – Careful procedures while satellite integration – RTGs are an interference source for plasma diagnostic  equipment, which has to be deployed away. – Space launch has a considerable  explosion risk. In this situation,  radioactive material can be dispersed  in the atmosphere.  [http://history.nasa.gov/computers/Ch6‐3.html] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 32 • Nuclear fission systems – Similar to ground nuclear reactors – Started with SNAP project in 60’s: RTG developed in SNAP project – Continued in 80’s with SP100 project • Many reactor models: with heat pipes transporting heat to a thermo‐ionic  reactor, cooled with lithium, with liquid metal with electromagnetic pumps… – Continued nowadays with SAFE project • Use of Brayton Cycle gas turbine, with uranium nitrium core, up to 100 kWe • Tested in 2001 in SAFE‐30 experiment (17.1 kW) The Role of Nuclear Power and Nuclear Propulsion in the Peaceful Exploration of Space, International Atomic Energy Agency http://web.archive.org/web/20050321055406/ http://www.spacetransportation.com/ast/presentations/7b_vandy.pdf
  86. 86. 17 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 33 • Solar heat systems – Original concept for the ISS – Use of solar energy to heat a fluid and move a rotary converter  (solar dynamics) or a thermoelectric converter (solar thermoelectric) – Conversion efficiency is 25% greater than for photovoltaic systems – Use of a Brayton Cycle engine – Solar light heats a fluid, helium mixed with xenon, no corrosion – Power storage thermally using latent heat of fusion for lithium  fluoride/calcium fluoride – Mass savings as compared  to battery technology [Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 34 • Secondary power systems – Batteries • Provide energy while primary system is not available  • Have to be recharged while first system is available  • Example: Batteries are used while solar arrays are in shadow,  and are recharged while under sunlight
  87. 87. 18 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 35 • Battery state of charge example – 9 am – 9 pm orbit [UPCSat‐1: Analysis, Design and Breadboarding of a University Picosatellite, A. Sánchez, J. Serra, A. Camps, M. Domínguez] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 36 • Choosing the right batteries – All batteries characteristics depend on materials • Cell voltage • Energy Density • Charge efficiency • Cycles at certain depth of discharge (DOD) • Radiation sensitivity • Temperature sensitivity Material Cell  Volts Capacity (W∙h/kg) Temperature  range Cycles @ 75%  DOD Cycles @ 25% DOD Ni‐Cd 1.25 30‐40 ‐20° ‐ 45° 800 21000 Ni‐H2 1.30 50‐80 0° ‐ 20° 4000 150000 Ag‐Zn 1.10 60‐70 0° ‐ 50° 100 3500 Ni‐MH 1.2 60 0° ‐ 25° 4000 130000 Li‐Ion 3.7 100‐250 ‐25° ‐ 75° 700 2700 [Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003]
  88. 88. 19 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 37 • System sizing  – While operative, primary energy system should be able to • Feed all systems • Recharge batteries more energy than they will spend – Batteries should be able to store more energy than which is going to  be consumed during eclipse – Both systems should be oversized because of degradation – We have to consider the worst situation T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 38 • Required power – Required power at EOL (End Of Life) depends on: • PLOADS: Power required by worst case in sunlight [W] • Pe: Power required by worst case in shadow [W] • te: Eclipse period [s] • ts: Sunlight period [s] • : Battery Charger efficiency (∼0,92) • : Battery Discharger (regulators, buses…) efficiency (∼0,9) • : Shunt dump efficiency 
  89. 89. 20 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 39 • Solar array sizing – Solar arrays must be sized depending on the required power at EOL: cos • S = Solar intensity (∼1400 W/m2) • cosφ= Angle between Sun line and array normal – If we are calculating for more than one array and/or geometry  is not planar, we should replace it by “Solar Panel Orientation  Coefficient” • = End Of Life cell conversion efficiency  • F = Sum of loss factors (∼0.75) • PF=Packaging efficiency  (∼0.85): Part of solar arrays made by solar  cells, taking into account void spaces, connections… • ASA=Solar Array Surface T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 40 • Solar panel orientation coefficient – Defines how many solar panels are illuminated pondering – Illumination time (% of time under light and in shadow) – Incidence angle – Number of panels – Example: Cubesat with all sides with solar arrays, rotating, with  one axis fixed towards the Earth’s magnetic field – Orbit 6 am – 6 pm (100% sun): 1,408 x 1U – Orbit 9 am – 9 pm (73% sun): 1,528 x 1U – Orbit 12 am – 12 pm (64% sun): 1,473 x 1U
  90. 90. 21 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 41 • Battery sizing – Battery capacity should be chosen according  to: • EBAT = Battery capacity [A∙h] • NB = Number of batteries • NC = Number of cells • VD = Cell discharge voltage [V] • DOD = Depth of discharge T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 42 • Power budget – Study of how much power is required by each subsystem as a  function of time, and how much power is available – Need to know • Load power consumption and on‐off time • Systems power consumption and on‐off time • Battery capacity • Available power • Light and eclipse time • Subsystems efficiency
  91. 91. 22 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 43 • Power budget example – UPCSat‐1 power budget [UPCSat‐1: Analysis, Design and Breadboarding of a University Picosatellite, A. Sánchez, J. Serra, A. Camps, M. Domínguez] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 44 • System sizing example (1/2) – A satellite consumes 6 W, in a 1 h of light and 15 min of shadow orbit • Batteries have to store more than: 6 W ∙ 15 min = 1.5 W∙h • Solar panels have to feed 6 W continuously 1,5 W∙h / 1h of sun = 1.5 W to recharge batteries 7,5 W in total – Neither batteries and solar panel degradation, nor system efficiency  has been taken into account  margings!
  92. 92. 23 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 45 • System sizing example (2/2) – If discharging system efficiency is about 90% • Total energy stored should be 1,5 W∙h / 0,9 = 1,66 W∙h – If planned mission time is about 3 years • 19,2 cycles/day = 21.000 cycles • If battery type is Ni‐Cd, at the end of the mission, DOD will be 25% • Battery should be able to store: 1,66 W∙h / 0,25 = 6,66 W∙h T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 46 • Power control and distribution – Array regulation: • Available and needed power vary during the mission • At the beginning of the mission, because of oversizing, available power is  higher than required • It is necessary to regulate power supplied by solar arrays – Sequential switching‐shunt regulation (S3R) – Pulse width modulation (PWM) and filters – Maximum peak power tracking (MPPT) – Battery control • Battery charge regulator • Battery management unit • Battery discharge regulator – Power control and distribution – Power regulation
  93. 93. 24 T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 47 • Battery charge regulator – Battery charging is a extremely delicate task – Each battery type has to be charged according to certain  parameters • Example: Li‐ion batteries have to be charged at constant  current until cell voltage reaches 4.2V. If cell voltage exceeds  that limit, it can explode. – Depending on charging current, battery useful life will change – NiCd battery reliability depending on overcharge factor (K) [Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003] T5. Electrical Power Supply System IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 48 • Battery management unit – Monitors battery states • Voltage, current • Temperature, pressure – Provides control inputs to BCR – Interface between data‐handling subsystem and EPS (senses telemetry) • Battery discharge regulator – Protects battery against high discharge currents that can damage them • High discharge currents reduce battery useful life • Really high discharge current can make explode batteries • Some high currents are caused by latch‐ups and SSE – Allow batteries to full discharge in order to improve battery capacity

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