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Team Members:Yuval PoratMoshe SederoHanan AmarNoam LeshemNoam LeiterOshrat MarfogelIttai CohenNitsan BavliSupervisor:     ...
Contents:1. Background                 8. Structure2. Requirements               9. ADCS sub-system3. Completed PDR Design...
Background• The oceanic surrounding is hazardous and present risks of  drowning , hypothermia, shark attacks and more…• Du...
Customer Requirements1. The system shall locate and a person in distress in any watery surrounding around   the world (oce...
Top Level Mission Requirements•    A user in distress shall be detected in less than 15 minutes, from signal transmission ...
Completed PDR Design Work:Electric Power System:  EPS + Matching        Solar Array to        Battery to Consumers        ...
GPSCommunication:                                                                                                         ...
Launch Segment:1. Poly-PicoSatellite Orbital Deployer “P-Pod MKIII”:•    mass 1.5 kg•    can carry 3 (1U) cubesats or 1 (3...
PDR Summary - Mission:Constellation : 700 km , i 45, e 0 , 48 satellites, 6 planesGeolocation: TDOA algorithm, 97% locatio...
PDR Summary - System:Mass: 3.11 kg                   Communication: 2 dipole, receiver, transmitterThermal Ctrl: Passive  ...
System Engineering                  (Updates)6 November 2011
Mission Profile                  ˆ                  x                                  ˆ                                  ...
Budgets      ΔV Budget                              Mass Budget                               Power Budget                ...
Design Iteration:Subsystem’s Mass:                                     Satellite’s Mass:                   Allocation for ...
System Hierarchy Diagram:6 November 2011
Satellite Block Diagram                                   2 UHFS-Band                                                   An...
Physical Hierarchy:                    2System interfaces ( N diagram):6 November 2011
Mission Design                     (Updates)6 November 2011
Orbits and ConstellationPDR Summary:• A Walker constellation 45:24/6/1• Constellation altitude - 700 km• Constellation inc...
Design Updates Since PDR:• Altitude had been changed from 700 km to 710 km.• At 710 km ionization dose is about 6 krad for...
De-Orbiting                                                                      m• PDR calculation for de-orbiting from 7...
• NASA’s de-orbit mechanism increases satellite’s surface area, and  thus drag force, by 60%• Device’s dimensions:        ...
Alternative #2: Tether Unlimited © nanoTerminator• Designed and manufactured be Tether Unlimited ©Specifications:• Mechani...
Method of Operation:• The conductive tape produces  current up the tape upon  interaction with ionspheric plasma• Charged ...
Device’s Performance:• The extended tape’s surface area is about 152 cm², increases  spacecraft surface area by 50 %• Deor...
De-Orbit Mechanism Selection            CriterionCriterion             Weight                         Propulsion Based    ...
GeolocationThe TDOA location methodt21      1        s2 u                     1                       s1 u                ...
•   If no measurement errors exist the target must lie on the hyperboloid defined    by this quadratic form where the 2 sa...
•   If the targets location is known to be constrained on the surface of the Earth only 2    more TDOA’s are needed to fin...
•   In the presence of measurement errors the initial location can be far from the true    location of the target. In orde...
Experiment Scale Down• In order to improve the reliability of the geolocation algorithms and  examine them in a more reali...
Acoustic TDOA Experiment:The satellite formation is hovering on a 4 on 4 meters air table. The target ismounted 3 meters a...
Satellite Formation Flight and Target Location on Table Plane                                           Nominal target ran...
Evolution of Location Error                                           Time SD is 15[ sec] Svs Location SD is 1[cm]        ...
Final Estimation Error is 8.32[cm] with 3 = 20.9 [cm]              120              100              80       [cm]        ...
Formation Flying PDR summary• In the first semester we selected the following principals:   – 2 satellites per formation  ...
CDR Revisions:1. No Altitude Maintenance  –     Satellites are allowed to lose altitude  –     Once correction is needed, ...
2. Statistical Analysis• In an attempt to reduce ∆V required, we performed a statistical  analysis of the actual scenarios...
Satellite Design                      (Updates)6 November 2011
StructurePDR SummaryA 3U cubesat has been chosen for the satellite`s structure.Inner Components PlacementGuidelines:• Maxi...
CDR UpdatesTwo alternatives of the 3U cubesat were considered:         Pumpkin skeleton                 ISIS skeleton6 Nov...
Structure Selection                  Criteria       Criteria   Pumpkin structure   Isis structure                         ...
The satellite`s structure6 November 2011
The Satellite`s Three Major Areas Bottom - Electronics   Middle – Propulsion System   Top – Communications6 November 2011
Exploded View6 November 2011
Analysis  A Finite Elements method is required – In order to reduce the  complexity of the geometric model a simplified mo...
Modal Analysis The boundary conditions are fixed support on all eight legs of the skeleton in order to simulate the satell...
Modal Analysis                          1st mode   2nd mode   3rd modeThe first 6 modes are:Mode     Frequency [Hz]  1    ...
Static AnalysisA 16g load was set in the longitudinal direction and a 2.75g was set inthe lateral direction.The results sh...
Static Analysis Results   Middle – Deformations   Top – Stress   Entire Satellite Deformation6 November 2011
Attitude Determination & Control SubsystemRequirements1. Spacecraft shall be 3 axis stabilized2. Spacecrafts long axis sha...
PDR Review:• Attitude control actuators: 3 magneto-torquers.• Attitude Determination: Magnetometer + Analog Sun Sensors  (...
Hardware Updates:                             PDR                        CDR                                            Ho...
Analog Sun Sensor DesignCurrent to sun angle of attack relation:             ˆ                                            ...
The Current-Sun AOA Relations:       I1     I max cos         3   sin   1       I2     I max cos         3   sin   2      ...
Attitude Determination Algorithm                                                             I• Computing Sun Vector and M...
Problem: During Eclipse sun’s Vector in body frame is unattainable.Consequence: Attitude determination of the satellite du...
Simulation ResultsSimulation time: 2 days. Disturbance Forces: STK Default 6 November 2011
Control Designthe control algorithm needs to deal with the following disturbances:• Gravity                               ...
Control Design – State-SpaceOur state-space equations will be:                                                          ...
Control Design – Control System Topography                                                  euler                         ...
Control Design – Results                                                         Steady State                             ...
6 November 2011
ΔV Budget                                      ΔV[m/s]                    Usage                                PDR        ...
PDR Summary• There were 3 missions for the Propulsion System:   1.   Positioning.   2.   Keeping formation.   3.   Deorbit...
Design updates since PDR• There are 2 missions for the Propulsion System:    1.   Positioning.    2.   Keeping formation.•...
The Propulsion Systems Block Diagram Block Diagram And Detailed Components                              Pressure    Straig...
Strength Analysis And Optimization• In order to design the most optimal  components, we made analysis with  “SimulationXpr...
Design Parameters OptimizationIn order to choose the most suitable design parameters we made graphs and atiterative way we...
Cold Gas Thruster - SpecificationsThe Cold Gas Thruster parameters:(T≅278K)Parameter          Value             Parameter ...
Programmatic Design6 November 2011
Cost Estimation - Propulsion example                                           Gas                 Pressure     Pressure  ...
Cost Estimation                                          Recurrent cost $       Components                                ...
Risk managementEvery project has risks –uncertainties that werent anticipated earlierRisk Management -identifying, analyzi...
Risks analysis-The 7 major risks in the project                                                  Likelihood Not Likely    ...
Risk                                Pf        C           R-Risk FactorPropulsion system: Technical risk-Center of mass   ...
Summary:                                               Number of risks                                        system  The ...
System ReliabilityReliability: “the probability that a device will work without failureover a specific time periods or amo...
For our mission t=2 years and is taken as constant,                                                     2 yearsso reliabil...
Mission ReliabilityIn order to calculate the mission reliability we calculated thereliability of each phase of the mission...
Summary - Compliance to Requirements:                                        Mission                           Requirement...
AcknowledgmentsWe’d like to express our appreciation and gratitude to allThose who have helped us:Prof. Pini Gurfil, Dr. D...
6 November 2011
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2011 06 17

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  • על מנת לפשט את המשדר ככל האפשר, אין תקשורת בין המטרה והלוויינים למעט פולסים בודדים שמשדרת המטרה החל מרגע מסוים. ללא מידע על זמן או מיקום השידור שיטת האיתור שנבחרה היא מדידת הפרש הזמנים בקליטת הפולסים. כל הפרש זמנים מתאר היפרבולואיד במרחב שזוג הלוויינים הם המוקדים שלו.
  • בהיעדר שגיאות מדידת זמנים ומיקומי לוויינים מיקום המטרה מאולץ על כל אחד מההיפרבולואידים המתאימים למדידות הפרשי הזמנים.באמצעות 3 מדידות ניתן להגדיר 3 היפרבולואידים, אשר בחיתוך שלהם ממוקמת המטרה.
  • אם ידוע כי המטרה ממוקמת על פני כדוה"א, ניתן להשתמש באילוץ נוסף זה יחד עם שתי מדידות הפרשי זמנים על מנת לאתר את המטרה.כאשר מבנה של 3 לוויינים קולט שידור אחד ומפיק ממנו שתי מדידות הפרש זמנים, קיים פתרון אנליטי למציאת מיקום המטרה המשדרת על פני כדוה"א.על מנת לצמצם את מספר הלוויינים במבנה למינימום הכרחי של 2 לוויינים, הרחבנו את השיטה האנליטית לשיטה איטרטיבית המאפשרת באמצעות קליטה של 2 פולסים לאתר את המטרה המשדרת על פני כדוה"א.
  • בנוכחות רעש מדידה, המיקום הראשוני המתקבל משתי מדידות באמצעות השיטה האיטרטיבית עשוי להיות לא מדויק מספיק על מנת לעמוד בדרישות המשימה.באמצעות מספר מדידות גדול ניתן לשפר את המיקום הראשוני באמצעות משערך קלמן מורחב. בדוגמא שלפנינו ניתן לראות תוצאות של סימולציה בה המיקום הראשוני התקבל מחוץ למעגל 1 ק"מ, ולכן חורג מדרישות המשימה.באמצעות מספר עשרות מדידות נוספות שגיאת המיקום מצטמצמת למספר עשרות מטרים בלבד.
  • על מנת לבחון את אלגוריתם האיתור בתנאים אמיתיים יותר, ולא רק בסימולציה ממוחשבת, ערכנו ניסוי במעבדה למערכות חלל מבוזרות, שבמכון אשר לחקר החלל.האתגר בניסוי זה הוא לבחון מערכת המתוכננת לטווחים של מאות ואלפי ק"מ במעבדה על פני כדוה"א. מכיוון שבמערכת המתוכננת לחלל הפולסים נעים במהירות האור,הפרש הזמנים שהיה נמדד במעבדה על פני מטרים בודדים בין שני מקלטים היה מספר ננו שניות. בהיעדר יכולת למדוד הפרשי זמנים כאלו, פולס השידור האלקטרומגנטי הוחלף בשידור אקוסטי כך שהפרשי הזמנים הנמדדים הם באותו סדר גודל של הזמנים המתוכננים למערכת בחלל.
  • לתאר את המעבדה ולהראות עוד תמונות.
  • ניתן לראות את מסלול שני הלוויינים על שולחן האוויר, ואת היטל המטרה על השולחן (המטרה ממוקמת כ- 3 מטרים מעל השולחן).
  • תוצאות שערוך מיקום המטרה באמצעות אותו האלגוריתם בו בוצעו הסימולציות למעט שינוי פרמטר מהירות הפולס ממהירות האור למהירות הקול.במקום אתחול לפי אילוץ מטרה לפני כדוה"א, אתחלנו את המיקום הראשוני לפי גובה המטרה מעל השולחן. סטיות התקן של מדידות הזמנים התקבלו כתוצר נוסף של תוצאות הניסוי.
  • כאן מוצג תקציב הדלתא וי למשימה. ניתן לראות, שעיקר תפקידה של מערכת ההנעה הוא למקם את הלוויין במסלולו. בנוסף לכך, משתמשים בהנעה גם לשמירת מבנה.
  • Transcript of "2011 06 17"

    1. 1. Team Members:Yuval PoratMoshe SederoHanan AmarNoam LeshemNoam LeiterOshrat MarfogelIttai CohenNitsan BavliSupervisor: :Jacob Herscovitz Winter 2010-2011
    2. 2. Contents:1. Background 8. Structure2. Requirements 9. ADCS sub-system3. Completed PDR Design 10. Propulsion sub-system and Summary 11. Cost Estimation4. System engineering 12. Risk Management5. Orbits and Constellation 13. Reliability6. Geo-location 14. Work Breakdown Structure (WBS)7. Formation Flying 15. Summary and Acknowledgments 6 November 2011
    3. 3. Background• The oceanic surrounding is hazardous and present risks of drowning , hypothermia, shark attacks and more…• Due to the nature and size of the oceanic surrounding, the process of receiving distress signals and locating people in distress accurately is somewhat problematic.• Between hundreds to thousands of sea-related accidents occur every year.6 November 2011
    4. 4. Customer Requirements1. The system shall locate and a person in distress in any watery surrounding around the world (oceans, seas, rivers…)2. The user shall wear an emergency beacon that will transmit a distress signal when activated.3. The time interval from distress signal transmission to notification in one of the ground stations shall not exceed 15 minutes.4. The computed location of the person in distress shall be no more than 1 km of his true location.5. As an option, the system shall allow enhanced capability for future applications such as search and rescue services for “land incidents”, given the appropriate modifications.6. The system shall be based on space and satellites technology.7. The space segment should be implemented using Nano-satellites ("Cube-Sat").8. Each satellites mission life-time shall be at least 2 years. 6 November 2011
    5. 5. Top Level Mission Requirements• A user in distress shall be detected in less than 15 minutes, from signal transmission to ground station notification.• A user in distress shall be geo-located with an accuracy < 1 km• The distress signal shall be relayed to a ground station• The systems services shall be affordable to the common end user.• The system shall be capable to identify its users in distress, as valid subscribers.• Earth coverage range shall be at least between latitudes +60⁰) and (-60⁰)• International space-related standards and regulations should be met, as much as possibleTop-level System Requirements:• "Cubesat" satellite platforms shall be considered.• Each satellites mass shall be less than 10 kg• Each satellite life time shall be at least 2 years.• Satellite bus shall be designed using space-proven COTS sub-systems and components, as much as possible.• Satellites sub-systems shall withstand launch load and space environment.• Geo-location shall be performed using DTOA technique, using 2 or 3 reception satellites. 6 November 2011
    6. 6. Completed PDR Design Work:Electric Power System: EPS + Matching Solar Array to Battery to Consumers Max Mass Battery Battery Efficiency Efficiency DoDClydeSpace 3U EPS + Battery Pack 90% 90% 20% 170 g Efficiency @ 5E14 e- Solar Panel Qty. Efficiency (BOL) Cell Area Cell Weight /cm2 Azure TJ 3G30C 26 29.1% 30.18 cm2 2.6 g 26.5%Thermal Control:•Passive Control•Steady State mean temperature: -6⁰ 6 November 2011
    7. 7. GPSCommunication: Other CubeSat User Geolocation Satellite Telemetry Main CubeSat Segment to Satellite & Control Satellite Ground Station Satellite Ground Station Antenna Monopole Patch Parabolic 2 Dipole 2 Dipole Yagi DipoleTransmitter 2.4Ghz 2.4Ghz -- 450Mhz 400Mhz 400Mhz Receiver -- 2.4Ghz 2.4Ghz 450Mhz 400Mhz 400Mhz 6 November 2011 Users Beacon Ground Station Ground Station for GeoLocation for Control andBroadcasting: data receiving and handling Telemetry6 November 2011
    8. 8. Launch Segment:1. Poly-PicoSatellite Orbital Deployer “P-Pod MKIII”:• mass 1.5 kg• can carry 3 (1U) cubesats or 1 (3U) cubesats• number of deployers can be mounted together on a L.V2. Launch Vehicle: SpaceX - Falcon 1e PayloadInclination Mass capability Est. Altitude space Accuracy Reliability [deg] [kg] Cost [m]Any above D1.55 x i = 0.1 [deg] LEO 800 to 700[km] Med $10.9M 9⁰ H1.7 Apogee = 15[km]6 November 2011
    9. 9. PDR Summary - Mission:Constellation : 700 km , i 45, e 0 , 48 satellites, 6 planesGeolocation: TDOA algorithm, 97% location within 15 min, 3% location within 30 minFormation: 2 satellites, In-plane formation, relative control, distance = 200 ± 50 km 6 November 2011
    10. 10. PDR Summary - System:Mass: 3.11 kg Communication: 2 dipole, receiver, transmitterThermal Ctrl: Passive Payload: Patch antenna, transceiverAttitude Ctrl: Active, 3-axis Propulsion: Warm gas, Isp=100Available Average Power: 6.78 W6 November 2011
    11. 11. System Engineering (Updates)6 November 2011
    12. 12. Mission Profile ˆ x ˆ y ˆ z Launch and Initial Dispersion Mission De-orbiting Deployment stabilization Operation 10 Mins 24 hours 14 days 2 years 1-2 years6 November 2011
    13. 13. Budgets ΔV Budget Mass Budget Power Budget PDR CDR Power consumption [mW] PDR CDR Sub System Consumers Usage System Total System Total ΔV[m/s] ΔV[m/s] Cruise Detection Maneuver Mass [Kg] Mass [Kg]Positioning OBDH 0.08 0.08 OBDH 200 600 600 Keeping ADCS 0.209 0.09 ADCS 430 630 630Formation Propulsion 1.209 0.458 Propulsion 0 0 2000Deorbiting 0 Thermal 0 0 Spare Thermal Control 0 0 0 0.94 Control (20%) Communication 200 450 450 Communication 0.23 0.23 Total 10.31 Payload 0.105 0.105 Payload 200 450 450 GPS 0.003 0.003 GPS 200 200 200 Power 0.297 0.237 EPS 200 200 200 Structure 0.958 1.02 Structure 0 0 0 De-Orbit - 0.08 De-Orbit 0 0 0 Total 3.111 2.3 Total 1430 1880 4530 6 November 2011
    14. 14. Design Iteration:Subsystem’s Mass: Satellite’s Mass: Allocation for Mass [Kg] Comments Sub-System Sub-System Sub-System Dry Mass 2.3 Total Mass [Kg] [Kg] 10% Margin 2.53 X+10% Power 0.31 0.2376 Includes 10% Margin for Fuel 0.031 ADCS 0.22 0.09 Fuel Thermal Control 0 0 Includes: Communication 0.265 0.23 Total 2.56 10% margin for Fuel and Payload 0.11 0.105 10% margin for Dry Mass GPS 0.004 0.003 OBDH 0.1 0.08 Propulsion 0.7 0.458 Structure 1.05 1.02 De-Orbit 0.2 0.08 Total 2.955 2.3 6 November 2011
    15. 15. System Hierarchy Diagram:6 November 2011
    16. 16. Satellite Block Diagram 2 UHFS-Band Antennas GPSAntenna S-Band Comm. OBDH UHF Comm. De-Orbit Device Antenna S-Band Main UHF UHF “Nano Receiver computer Transmitter Receiver Terminator” GPS Power Attitude Determination & Control Propulsion EPS Pr. Tank 3x 3x Magneto- Magneto- Battery Torquers Meters Valves Filter Photo Voltaic Cells Regulator 2x Thruster 6 November 2011
    17. 17. Physical Hierarchy: 2System interfaces ( N diagram):6 November 2011
    18. 18. Mission Design (Updates)6 November 2011
    19. 19. Orbits and ConstellationPDR Summary:• A Walker constellation 45:24/6/1• Constellation altitude - 700 km• Constellation inclination of 45⁰• Total of 48 satellites.• Total of 24 formations• 2 satellites per formation with nominal distance of 200 km between satellites.• 6 orbital planes, each orbital plane consisting of 8 satellites (4 formations)• Satellite de-orbitization at EOL using propulsion to lower the satellite from 700 km to 650 km, requiring v 13.35 sec m6 November 2011
    20. 20. Design Updates Since PDR:• Altitude had been changed from 700 km to 710 km.• At 710 km ionization dose is about 6 krad for 0.6 mm shielding thickness.  still well within the 10 krad restriction of the sensitive EPS system.• In 2 years (mission life time) satellites decline approximately 10 km.  at EOL, altitude is around 700 - higher than the minimum of 697 km• No altitude correction maneuvers are required throughout the entire mission. Constellation Revisit Time Vs Altitude Revisit Time [min] Altitude [km] 6 November 2011
    21. 21. De-Orbiting m• PDR calculation for de-orbiting from 700 km to 650 km: v 13.35 sec• From 710 km to 650 km – even higher: v 16.01 m sec• 2 Alternatives for De-Orbiting were considered:Alternative #1: “Jack in the Box”• De-Orbit mechanism designed and manufactured by NASA for the O/OREOS mission.6 November 2011
    22. 22. • NASA’s de-orbit mechanism increases satellite’s surface area, and thus drag force, by 60%• Device’s dimensions: Material: Aluminum plates Germanium Film 28 cm 9.9 cm 9.9 cm Weight: ~200 gr (est.) Device can be placed only on top or bottom panel6 November 2011
    23. 23. Alternative #2: Tether Unlimited © nanoTerminator• Designed and manufactured be Tether Unlimited ©Specifications:• Mechanism consists of a 30-m long, 0.8-mm thick conductive tape.• Mechanism can be mounted on every panel.• Photo-voltaic sells can be integrated onto it• Mechanism mass is ~ 80 grams6 November 2011
    24. 24. Method of Operation:• The conductive tape produces current up the tape upon interaction with ionspheric plasma• Charged tape interacts back with earth’s magnetic field to produce Lorentz Force that opposes orbital motion and produces electrodynamic drag.  L   F I B dl 06 November 2011
    25. 25. Device’s Performance:• The extended tape’s surface area is about 152 cm², increases spacecraft surface area by 50 %• Deorbit Time Prediction with mechanism: CAESAR Satellites6 November 2011
    26. 26. De-Orbit Mechanism Selection CriterionCriterion Weight Propulsion Based "Jack-in-the-Box" nanoTerminator Value Score Total Value Score Total Value Score Total Mass 0.5 400 gr 1 0.5 200 gr 3 1.5 80 gr 5 2.5Deorbit 0.2 24.4 yr 2 0.4 22.2 yr 3 0.6 <1 yr 5 1 TimeCompat- 0.3 1 0.3 3 0.9 5 1.5 ibility Total 1.2 3 5The nanoTerminator gives us the best deorbit time, for the lowestadditional mass, and is the easiest to integrate with the satellite. 6 November 2011
    27. 27. GeolocationThe TDOA location methodt21 1 s2 u 1 s1 u • The hyperbolic equation c c can be transformed to a 2 2 2 si u Xi X Yi Y Zi Z quadratic formc is the speed of light M m u uT Mu 2mT u m0 0 uT 1 0 mT m0 1 T M 4 s1 s2 s1 s2 4d 2 I m 2 2 s2 s12 s1 s2 2d 2 s1 s2 2 2 m0 s2 s12 d2 2 s12 2 s2 d2 d c t216 November 2011
    28. 28. • If no measurement errors exist the target must lie on the hyperboloid defined by this quadratic form where the 2 satellites in the formation are the focal of the hyperboloid. In this case 3 TDOA measurements can define the 3 unknown target coordinates. Satellite Formation and Target on TDOA Hyperboloid Satellite Formation and Target on TDOA Hyperboloid SAT1 SAT1 SAT2 SAT2 Target Target TDOA Hyperboloid TDOA Hyperboloid 4 3 2 -4 1 -3 0 -2 Z -1 -1 -2 0 -3 5 1 -4 2 4 5 0 3 2 4 Y 0 0 -2 -5 5 -4 -5 X Y X6 November 2011
    29. 29. • If the targets location is known to be constrained on the surface of the Earth only 2 more TDOA’s are needed to find the location.• Based on the analytical solution shown by Ho and Chan for a 3 satellite formation and a single TDOA measurement, we have derived an iterative algebraic method for a 2 satellite formation using 2 TDOA measurements. Target in the Intersection of a Sphere and 2 TDOA Hyperboloids SAT11 SAT21 Target SAT12 SAT22 8 6 4 Z 2 0 -2 5 5 0 0 -5 -5 -10 Y X6 November 2011
    30. 30. • In the presence of measurement errors the initial location can be far from the true location of the target. In order to improve the initial location error an Extended Kalman Filter starting with the initial location is used with all of the TDOA data. The estimated target location then drifts from the initial location closer to the true location.6 November 2011
    31. 31. Experiment Scale Down• In order to improve the reliability of the geolocation algorithms and examine them in a more realistic environment we have conducted an experiment at the Distributed Space Systems Laboratory (DSSL) in the Asher Space Research Institute. Parameter Space Scale EchoLab Scale Formation ~200 km ~500 mm Target Range 700-3000 km 3000-4000 mm V phase EM 300e3 km/sec Acoustic 340e3mm/sec TDOA 0-300 microsecond 0-300 microsecond SD time ~50 nanosecond ~50 microsecond SD length ~0.015 km ~17 mm6 November 2011
    32. 32. Acoustic TDOA Experiment:The satellite formation is hovering on a 4 on 4 meters air table. The target ismounted 3 meters above the table and transmits 40 KHz acoustic pulses. Satellites Ultrasound Transmitter 6 November 2011
    33. 33. Satellite Formation Flight and Target Location on Table Plane Nominal target range is 3.089[m] Nominal distance in formation is 0.617[m] 0.2 0.1 0 -0.1 -0.2 Target Y [m] SAT1 -0.3 SAT2 -0.4 -0.5 -0.6 -0.7 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 X [m]6 November 2011
    34. 34. Evolution of Location Error Time SD is 15[ sec] Svs Location SD is 1[cm] 0.2 3 0 Z [m] 2.95 -0.2 0.1 0 -0.4 -0.1 -0.6 -0.2 X [m] Estimation -0.8 Y [m] Initial Target6 November 2011
    35. 35. Final Estimation Error is 8.32[cm] with 3 = 20.9 [cm] 120 100 80 [cm] 60 40 20 0 2 4 6 8 10 12 14 16 18 20 22 [Estimation Steps]6 November 2011
    36. 36. Formation Flying PDR summary• In the first semester we selected the following principals: – 2 satellites per formation – In-plane formation – Relative control method Nominal State Distance Reaches Elliptic Verify Successful + Altitude Boundary Maneuver Maneuver Maintenance• This means that in worst-case scenario, ∆V required to m maintain formation and altitude is V = 7.774 sec 6 November 2011
    37. 37. CDR Revisions:1. No Altitude Maintenance – Satellites are allowed to lose altitude – Once correction is needed, the maneuvering satellite also changes its altitude to that of its partner’s (Hohmann Transfer) – In worst-case scenario, ∆V required to maintain formation is: m V = 2.52 sec Elliptic Distance Maneuver Reaches Verify Successful Nominal State Boundary Hohmann Maneuver Transfer6 November 2011
    38. 38. 2. Statistical Analysis• In an attempt to reduce ∆V required, we performed a statistical analysis of the actual scenarios that may occur. – Euler angles of satellite are normally distributed ( 0, 2.5 ) – 800 simulations performed• Results: – 69% of cases – No correction required – 31% of cases – 1 correction required – In none of the cases were 2 corrections needed m• Conclusion: ∆V needed to cover 99.99% of all cases is V = 0.18 sec6 November 2011
    39. 39. Satellite Design (Updates)6 November 2011
    40. 40. StructurePDR SummaryA 3U cubesat has been chosen for the satellite`s structure.Inner Components PlacementGuidelines:• Maximum distance between magnetometer and magnetic field generators (magnetorquers, electrical components)• Center of mass should be as close to geometric center as possible• Thrust vectors should pass as close to the center of mass as possible• Patch antenna facing Nadir direction• GPS Antenna facing Zenith direction6 November 2011
    41. 41. CDR UpdatesTwo alternatives of the 3U cubesat were considered: Pumpkin skeleton ISIS skeleton6 November 2011
    42. 42. Structure Selection Criteria Criteria Pumpkin structure Isis structure weight Compatibility 0.5 8 10 to ISIS ISIPOD Flight heritage 0.1 8 1 Modularity 0.4 7 9 Total 7.6 8.7 The Isis structure was chosen.6 November 2011
    43. 43. The satellite`s structure6 November 2011
    44. 44. The Satellite`s Three Major Areas Bottom - Electronics Middle – Propulsion System Top – Communications6 November 2011
    45. 45. Exploded View6 November 2011
    46. 46. Analysis A Finite Elements method is required – In order to reduce the complexity of the geometric model a simplified model was suggested:• The inner components are referred as “Point Mass”• Three mass points simulate the three major parts• Each point mass is connected through 8 points to the satellite’s skeleton in order to simulate the real assembly 6 November 2011
    47. 47. Modal Analysis The boundary conditions are fixed support on all eight legs of the skeleton in order to simulate the satellite in the launch POD.6 November 2011
    48. 48. Modal Analysis 1st mode 2nd mode 3rd modeThe first 6 modes are:Mode Frequency [Hz] 1 695 2 708.11 3 755.18 4th mode 5th mode 6th mode 4 756.94 5 769.25 6 769.6 6 November 2011
    49. 49. Static AnalysisA 16g load was set in the longitudinal direction and a 2.75g was set inthe lateral direction.The results show the satellite will endure the launch loads even with a10 degree misalignment with its long axis. 16g 10 deg6 November 2011
    50. 50. Static Analysis Results Middle – Deformations Top – Stress Entire Satellite Deformation6 November 2011
    51. 51. Attitude Determination & Control SubsystemRequirements1. Spacecraft shall be 3 axis stabilized2. Spacecrafts long axis shall be Nadir Oriented3. Maximum pointing error (per axis): 1. Cruise Mode: less than 5⁰ 2. Engine Ignition: less than 10⁰4. ADCS sub-systems mass shall be less than 190 grams5. Maximum power consumption shall be less than 630 mWatt6. Maximum time from deployment from launch pad until initial stabilization shall be less than 24 hours 6 November 2011
    52. 52. PDR Review:• Attitude control actuators: 3 magneto-torquers.• Attitude Determination: Magnetometer + Analog Sun Sensors (preliminary design)• Hardware Selection• Disturbance torque estimation 6 November 2011
    53. 53. Hardware Updates: PDR CDR Honeywell Billingsley HMC 5843Magneto TFM65-VQS (Integrated to OBC) -meter 117 gr 50 milligram 3.51x3.23x8.26 [cm³] 4x4x1.3 mm Satellite Services LTD Visio Torquer Torquer rod (x3) PCBMagneto-Torquer m = 30 gr m = 100 gr L=7 cm Size: 10 x 9 cm D=0.9 cm Dipole = 0.5 Am² Dipole=0.2 Am² 6 November 2011
    54. 54. Analog Sun Sensor DesignCurrent to sun angle of attack relation: ˆ y I I max sin ˆ x ˆ z I max is the current measured when the sun shines directly in the normal direction:  906 November 2011
    55. 55. The Current-Sun AOA Relations: I1 I max cos 3 sin 1 I2 I max cos 3 sin 2 CAESAR CAESAR I3 I max sin 3 Top View Side View I1Finding the AOA angles using: sin 2 cos 1 and 1 arctan I2Sun’s vector in Body Frame is written as: sin 1 cos 3 X 0 0 I1 VsB sin 2 cos 3  VsB 0 Y 0 I2 where, X , Y , Z 1 sin 3 0 0 Z I3 6 November 2011
    56. 56. Attitude Determination Algorithm I• Computing Sun Vector and Magnetic Vector in ECI - Vsun , Vmag I• Using sensor’s data to derive Sun Vector and Magnetic Vector in B B body frame: Vsun ,Vmag• Finding a rotation matrix from Body Frame to ECI: I I I I I B B B B 1 C B V sun V mag V sun V mag V sun V mag V sun V mag• Finally, Finding rotation matrix from Body Frame to VVLH: VVLH CB CIVVLH CB I• From the rotation matrix it’s easy to derive Euler angles by: C2,3 C1,3 C1,2 arctan , arctan , arctan C3,3 2 C 2,3 C 2 3,3 C1,16 November 2011
    57. 57. Problem: During Eclipse sun’s Vector in body frame is unattainable.Consequence: Attitude determination of the satellite during eclipse is unattainable.Solution: Rotational rate estimation from 3 attitude measurements, using Lagrange interpolation formula:  t t3 t 2 t3 t1 2t3 t1 t2 3 t1 t2 t3 t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2  t t3 t 2 t3 t1 2t3 t1 t2 3 t1 t2 t3 t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2 t3 t 2 t3 t1 2t3 t1 t2  t3 t1 t2 t3 t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t26 November 2011
    58. 58. Simulation ResultsSimulation time: 2 days. Disturbance Forces: STK Default 6 November 2011
    59. 59. Control Designthe control algorithm needs to deal with the following disturbances:• Gravity • Engine Torque• Solar Pressure• Atmospheric Drag yˆ• Magnetic Field ˆ x ˆ z g 6 November 2011
    60. 60. Control Design – State-SpaceOur state-space equations will be:          P Q R T A  I m b ng nd 0 0 0 0 0 0 0 0 0 0 1 0 0 0 0 0 0 0 0 0 0 0 0 1 0 0 0 0 0 0 0 m1 0 0 0 0 0 1 b3 b2 1  0 m2 0 0 nd 2 Ix Ix IxP 4 0 1 0 0 0 0 0 1 1 P b3 b1 m3 1Q 0 3 2 0 0 0 0 Q 0 0 0 0 2 Iy Iy IyR 0 0 2 2 1 0 0 R 0 3 0 3 b2 b1 0 0 1 0 Iz Iz IzWhile: ng – The gravity gradient disturbance moment nd – The remain disturbances moments m – The control dipole moment b – The magnetic field 6 November 2011
    61. 61. Control Design – Control System Topography euler distubances moments nd nd+ngg -K- Y=eye*X T u=-Kx r2d angles T_ctrl X1 err u uIn1 -K- StateSpace PID+ Anti WindUp r2d1 rates -K- Anti WindUp Ks PID -K- 1  s Integrator1 bin x Ki Saturation1 P 1 err -K- 1 u Kp Q 2 -K- rates R Kd Scope3 6 November 2011
    62. 62. Control Design – Results Steady State Eclipse Response 3 100 2 Angle [deg] 50 Angle [deg] 1 0 0 -1 0 1000 2000 3000 4000 5000 6000 7000 8000 time [sec] -50 0 1000 2000 3000 4000 5000 6000 7000 8000 time [sec] 180o command: Controller ST=4652.8566sec 250 200 150 Angle [deg] 100 50 0 -50 0 1000 2000 3000 4000 5000 6000 7000 8000 time [sec] 6 November 2011
    63. 63. 6 November 2011
    64. 64. ΔV Budget ΔV[m/s] Usage PDR CDR Positioning Keeping Formation Deorbiting Spare (10%) Total6 November 2011
    65. 65. PDR Summary• There were 3 missions for the Propulsion System: 1. Positioning. 2. Keeping formation. 3. Deorbiting.• We selected a warm gas propulsion system of “MicroSpace”.• We designed an external high pressure gas tank for the propulsion system.• The total propulsion system mass was 436 g.• The cost of the propulsion system without the external gas tank was € 81,000.6 November 2011
    66. 66. Design updates since PDR• There are 2 missions for the Propulsion System: 1. Positioning. 2. Keeping formation.• We noticed that “MicroSpace’s” propulsion system is too heavy, complicated and expensive. So, we designed a new Cold Gas Propulsion System that meet our specific requirements.• The total propulsion system mass is 429 g.• The cost of the propulsion system is $ 7,321.6 November 2011
    67. 67. The Propulsion Systems Block Diagram Block Diagram And Detailed Components Pressure Straight Main Transducer Pipe Connector Pressure Connector Pressure Solenoid Regulator Valve Fill Pressure Valve Regulator Straight High Latch Connector Pressure Valve Tank Curve Solenoid Pipe Pressure Valve Transducer Pressure Solenoid Regulator Valve Thruster Fill Valve House Latch Thruster Valve Gas Tank 6 November 2011
    68. 68. Strength Analysis And Optimization• In order to design the most optimal components, we made analysis with “SimulationXpress”.• At iterative work, we fit the wall thickness to the applied pressure(at extreme conditions of 50°C), so we get the optimal weight.6 November 2011
    69. 69. Design Parameters OptimizationIn order to choose the most suitable design parameters we made graphs and atiterative way we gathered to the best solution. Thrust Vs. Pc Thrust Vs. Area Ratio 0.8 tpulse Vs. Area Ratio 0.08 Thrust Vs. Pc Thrust Vs. Throat Diameter Thrust Vs. Area Ratio 0.6 0.08 0.8 420 1.5 0.075 F [N] F [N] 0.4 0.6 0.07 0.2 t pulse [sec] 0.0751 400 F [N] F [N] 0 0.065 0.4 F [N] 5 10 15 20 25 30 35 40 45 50 55 60 20 40 60 80 100 120 140 160 180 Pc [atm] Area Ratio Ae/At t pulse Vs. Pc Isp Vs. Area Ratio 0.5 380 80 3000 0.07 0.2 75 tpulse [sec] Isp [sec] 2000 0 0 1000 360 0.065 0.1 5 0.2 10 0.3 15 0.4 20 0.5 25 0.6 30 0.7 350.8 40 0.9 45 1 50 1.155 70 60 20 20 40 40 60 60 80 100[atm]120 140 160 180 Throat Diameter [mm] 140 80 100 Pc 120 160 180 Area Ratio Ae/At tpulse Vs.tRatio Vs. Pc 60 Area Ae/At 0 65 pulse Diameter Throat 5 10 15 20 25 30 35 40 45 50 55 20 40 60 80 100 120 140 160 180 Pc [atm] mIsp Vs. Area Ratio prop Vs. Area Ratio Area Ratio Ae/At 3000 3000 80 44 Thrust Vs. Throat Diameter tpulse Vs. Area Ratio 1.5 420 tpulse [sec] 2000 [sec] 1 2000 42 75 t pulse [sec] mtprop [gr] 400 Isp [sec] F [N] pulse 0.5 40 1000 380 1000 0 0.1 0.2 0.3 70 0.5 0.4 0.6 0.7 0.8 0.9 1 1.1 360 20 40 60 80 100 120 140 160 180 38 Throat Diameter [mm] Area Ratio Ae/At 0 0 tpulse Vs. Throat Diameter mprop Vs. Area Ratio 3000 5 10 15 20 0.5 25 0.6 30 0.7 350.8 40 0.9 45 1 50 1.155 60 36 65 0.1 0.2 0.3 0.4 44 20 20 40 40 60 60 80 100[atm] Pc 120 Throat Diameter 120 80 100 140 [mm] 140 160 160 180 180 tpulse [sec] 2000 42 mprop [gr] Area Ratio Ae/At Area Ratio Ae/At 40 1000 386 November 2011 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 Throat Diameter [mm] 0.9 1 1.1 36 20 40 60 80 100 120 140 160 180 Area Ratio Ae/At
    70. 70. Cold Gas Thruster - SpecificationsThe Cold Gas Thruster parameters:(T≅278K)Parameter Value Parameter ValueThroat diameter 0.3 mm 10.31 m/secExit diameter 3 mm Pulse time 370 sec Propellant mass (N2) 37.6 grThrust 75.2 mN Tank Pressure 137 atm = 2,015 PsiIsp 74.8 secPc 6 atmPe ~0 atm 6 November 2011
    71. 71. Programmatic Design6 November 2011
    72. 72. Cost Estimation - Propulsion example Gas Pressure Pressure Latch Solenoid Control Part Name Thruster Fill Valve Fasteners Tank Regulator Transducer Valve valve Board Parts Cost $ 303 66 1,500 885 900 233 400 452 724 Time [Min] 120 Rate Assembly Work 100 [$/Hour] Cost $ 200 Opacity Test - Time [Min] 120 Helium mass Rate Work 70 spectrometer [$/Hour] with bell jar Cost $ 140 Quantity for the constellation 48 96 96 48 48 96 48 48 48 Total Cost per Satellite $ 7,268 Total Cost for Entire Constellation $ 348,856 Pressing pattern $ 10,000 Non- Opacity Equipment $ 5,000 recurrent Environmental 30,000 testing $ Total non-recurrent cost $ 45,0006 November 2011
    73. 73. Cost Estimation Recurrent cost $ Components Cost for Entire Non-recurrent cost $ Cost per satellite Constellation Satellite Structure 6,585 316,113 Propulsion System 7,268 348,856 45,000 ADCS 19,604 941,008 Payload System 13,355 641,072 Communication System 23,149 1,111,192 Power 17,368 833,664 Formation 16,600 Geolocation 35,480 Total cost 87,329 4,191,905 97,080 Constellation cost 4,288,985 Launching the entire constellation (6 launches) costs : ~ $ 67,714,4646 November 2011
    74. 74. Risk managementEvery project has risks –uncertainties that werent anticipated earlierRisk Management -identifying, analyzing and responding to project risk.project risks are uncertainties that may result in schedule delays, cost overruns,performance problems, adverse environmental impacts or other undesiredimpacts.Pf - The likelihood of the eventC - The potential consequence to the projectR - Risk factor R P C f6 November 2011
    75. 75. Risks analysis-The 7 major risks in the project Likelihood Not Likely Likely Very Likely Consequence Low Risk Low Risk Low Risk Benign Low Risk Medium Risk Medium Risk Medium Low Risk Medium Risk High Risk Harsh Risk Pf C R-Risk FactorPropulsion system: Safety risk-the system contains 0.8 0.8high pressure, chance of explosion. 0.64Risk Mitigation: Performing experiments and tests on the system and particularly on the tank.Propulsion system: Schedule risk- the launch 0.8 0.7 0.56company will not agree to launch the satellite.Risk mitigation: Experiments and higher safety factors.Propulsion system: Technical risk- The amount of 0.7 0.7 0.49gas might not be enough -sun storms increase drag.Risk mitigation: Increasing the percentage of spare gas in the tank. This spare gas will be used inunexpected weather in space. 6 November 2011
    76. 76. Risk Pf C R-Risk FactorPropulsion system: Technical risk-Center of mass 0.7 0.6wouldnt coincide with the engine`s nozzle. 0.42Risk mitigation: Designing a new propulsion system or moving components in the satellite.Launch: Finding time windows suitable for the launch of 0.5 0.924 pairs of satellites in 6 different launch dates. 0.45Risk mitigation: Communicating with launch provider in advance as possible in order to decrease theprobability of such failure.Orbits and constellation: Technical risk-Satellite collision 0.4 0.9with space debris 0.36Risk mitigation: Running Debris assessment simulations using NASAs Debris Assessment Tool, andSTK. Launching redundant (extra) satellites to account for damaged satellites.Attitude control: Stabilization of the satellite by the 0.5 0.8attitude control system. 0.4Risk Mitigation: Testing the satellite in a laboratory and performing simulations. 6 November 2011
    77. 77. Summary: Number of risks system The propulsion system has the found most risks in the project and Propulsion 8 the consequence of its risks is Attitude control 2 the most severe. This is Geolocation 1 understandable since the Structure 2 launch 2 propulsion system is new and Formation 2 we don’t have previous Keeping experience with such systems. Orbits and 2 This means we would have to constellation Electrical Power 4 perform more experiments Communication 2 and tests on the system and of /payload the system with the satellite Thermo control 1 in order to mitigate the risks. Total6 November 2011
    78. 78. System ReliabilityReliability: “the probability that a device will work without failureover a specific time periods or amount of usage” *IEEE, 1984]. R e tR - Success Probability, - Failure Rate , t -Time PeriodSeries Reliability: A B C RS RA RB RCParallel/Redundant Reliability : A B RP 1 1 R A 1 RB 1 RC C 6 November 2011
    79. 79. For our mission t=2 years and is taken as constant, 2 yearsso reliability is computed as: R e dt t 0Example: Propulsion Subsystem Reliability R5=0.9801 R6=0.992 R7=0.99 Pressure Solenoid Thruster Regulator Valve Pressure Propellant Latch Fill ValveTransducer Tank Valve R1=0.988 R2=0.9999 R3=0.996 R4=0.9994 Pressure Solenoid Thruster Regulator Valve R5=0.9801 R6=0.992 R7=0.99 2Rpropulsion R1 R2 R3 R4 1 1 R5 R6 R7 0.96385466 November 2011
    80. 80. Mission ReliabilityIn order to calculate the mission reliability we calculated thereliability of each phase of the mission: Satellite Initial Mission RLLaunch 0.97 R 0.94 Stabilization S RPositioning Ph 0.91 ROperation M 0.902 RDDeorbit 0.97 (Phasing) De-Orbit ADCS ADCS Computer ADCS Computer Computer EPS EPS EPS Communication Communication Communication Communication EPS Propulsion Propulsion Payload Mechanism RMission RL RS RPh RM RD 0.735 6 November 2011
    81. 81. Summary - Compliance to Requirements: Mission Requirement Result Compliance Constellation Revisit Time < 15 min 14.74 min  Geo-Location Location Radius < 1 km 97% < 1 km  De-Orbiting within 25 years 1-2 years  Orbits Global Coverage between Available Coverage: latitudes +60⁰) and (-60⁰) +60⁰) and (-60⁰)  ~$87K per Sat Cost Cost-efficient ~$4.2M Total  System Satellite’s Mass Each Satellite’s mass < 10 kg 2.56 kg 6 November 2011
    82. 82. AcknowledgmentsWe’d like to express our appreciation and gratitude to allThose who have helped us:Prof. Pini Gurfil, Dr. David Mishne, Dr. Zvi Hominer,Dr. Avi Vershavski, Ofer Slama. And special thanks to our supervisor Jacob Herscovitz6 November 2011
    83. 83. 6 November 2011
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