MICROSCOPE : A Microsatellite for Equivalence Principle Measurement in Space V. CIPOLLA, J.B. DUBOIS, B. POUILLOUX, P. PRIEUR Centre National d’Etudes Spatiales 18 av. Edouard Belin Toulouse France Planning of presentation ■Scientific objectives ■Mission overview ■Payload description ■Satellite description ■Satellite performances ■Conclusion XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 2 Scientific objective ■(Weak) Equivalence Principle a.k.a. universality of free fall XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 3 Scientific objective ■Equivalence Principle (EP) definition Mass could be defined in two differential ways : • Inertial mass mi : proportional term between the force acting on a body and its accelerationr r = ⋅γ F mi • Gravitational mass mg : proportional term between the gravity forces acting on a body submitted to a gravitation field and the gravitation field itself m ⋅ M G ⋅ M = ⋅ g g = ⋅ g = ⋅ Fg G mg mg g r 2 r 2 Both masses are “identical” (i.e. different bodies have the same ratio) m m 1g = 2g m1i m2i XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 4 Scientific objective ■Interest of equivalent principle measurement EP is a fundamental assumption for the theory of general relativity EP has not theoretical justification however it has been always verified by experiments Violation of EP is predicted by several quantum gravity theories ■Microscope aims at testing EP with an accuracy better than 10 -15 : If EP is violated the limit of the validation of relativity theory will be identified If EP is confirmed some quantum gravity theories could be swept away XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 5 Mission overview ■Equivalence principle test in space To simulate free fall putting two different masses in a circular orbit around the Earth with the same initial conditions If PE is verified m m 1g = 2g m1i m2i If PE is violated m m 1g ≠ 2g m1i m2i Trajectory measurement is replaced by acceleration measurement (i.e. the force necessary to keep the CoG of the masses at the same location) m m Γ = 1 ⋅ ()γ − γ = 1 ⋅ 1g − 2g ⋅ = 1 ⋅δ ⋅ d 1 2 g g 2 2 m1i m2i 2 The accuracy of the experimental protocol is checked by a 2 nd differential accelerometer built with two proof masses made of an identical material XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 6 Mission overview ■Consequences of the accuracy goal on the mission Violation measurement in continuous mode is not achievable need to “modulate” the gravity field seen by the proof masses at the frequency ƒEP and use signal process method for data analysis • Inertial attitude of the S/C ƒEP = orbital frequency ƒOR • S/C rotating around the direction to the orbital plane at the frequency ƒSP ƒEP = ƒOR + ƒSP Measurement performed on several orbits to reduce noise effects Tolerance on PL manufacturing drive to a little difference on location of the CoG of the proof masses which perturb EP measurement • Need of specific calibration session • Need to minimize non-gravity forces acting on the S/C • Need of an accurate orbit determination on ground Guarantee an ultra stable environment to the PL • Minimization of all possible perturbation acting at ƒEP • Minimization of µperturbation (MLI thermoelastic clanks, moving masses, etc..). XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 7 Mission overview ■Main mission requirement Orbital parameters • Eccentricity < 0.005 EoL • Orbit position knowledge < 7 m at ƒEP Drag free requirement • -12 2 S/C acceleration induced by non gravity forces shall be < 10 m/s at ƒEP Attitude control requirement • Angular velocity stability < 10 -9 rad/s • Angular acceleration < 10 -11 rad/s 2 Thermal requirement • Thermal stability of Sensor Unit < 2 mK at ƒEP • Thermal stability of Payloads front end units < 20 mK at ƒEP XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 8 Payload Description ■T-SAGE : Twin Space Acceleration for Gravity Experiment Designed, developed and integrated by ONERA DMPH Châtillon (Fr) Composed by • 2 Sensor Units (SU) T-SAGE FEEU1 RS422 UART – Each SU composed by 2 concentric 1.25 Mbits electrostatic 6-axis inertial sensors SU1 OS-Link – SU 1 : different materials (Pt, Ti) – SU 2 ; same materials (Pt) SU2 • 2 Front End Electronics Units Secondary Voltage ±48V, ±15V, +5V – Capacitive detection Secondary Voltage – DAC ±48V, ±15V, +5V • I/F Control Unit ICUME SUMI ICUME BNR1N – SUs control loops BNR1R – I/F with the S/C BNR2N RS422 UART BNR2R 1.25 Mbits FEEU2 XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 9 Satellite description ■Microscope developed in the Frame of Myriade product line ■Myriade microsatellite Myriade development started in 1999 under lead of CNES • Target :150 Kg, 200 W and 2 years class micro satellite Design based on a Platform with generic functional chain : • Structure (customizable) • Power Payload • AOCS XSAT • On board data management ZSAT YSAT • Communication • Propulsion (option) Platform Generic multi-mission ground segment 11 satellites launched cumulating more than 35 years of orbit life 8 satellites ready to launch or in development XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 10 Satellite description ■Microscope specificity vs Myriade product line standard bus PL inside the S/C Structure specific development Drag free Attitude and Acceleration Control System • Allow to control S/C background acceleration and attitude stability • AACS specific equipments : – T-SAGE used as sensor unit (hybridization with Star Tracker measurements for attitude) – Cold Gas Propulsion System as actuator (Microscope specific development) French Space Operation Law Passive desorbitation system • No chemical propulsion available on board (to avoid slashing of propellant) • Deployment of two sails using Gossamer arms (increase of S/C Surface/Mass ratio) ■Other functional chain are identical to Myriade ones Power (solar generator geometry adapted) AOCS (for non scientific modes) On board data management (On board software to be adapted) Communication XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 11 Satellite description XSAT CGPS Y SAT Tank Payload Battery Module ZSAT RW PCDU ICUME OBC GNSS XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 12 Satellite description ■Microscope structure PF structure adapted from standard Myriade definition • 6 rectangular sandwich panels with Aluminium alloy skins and honeycomb core • Manufacturing procedure and integration principle issued from previous programs Payload structure (PAS) specific development • Composed by two stage metallic structure with its autonomous thermal control – 1st stage dedicated to SU – 2nd stage dedicated to FEEU • Design goals – Allows the centering of SU wrt satellite spin axis – Offer very high mechanical and thermal stability – Protect SU from external perturbations – Guarantee thermal control of SU and FEEU • Thermal control qualification test performed in 2009 – Stability of SU Temperature at ƒEP measured < 1 mK – Stability of FEEU Temperature at ƒEP measured < 17 mK XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 13 Satellite description ■Attitude and Acceleration Control System Based on Myriade standard AOCS for servitude modes : MLT, MAS, MGT Transition mode MSP and mission mode MCA specific to Microscope • MCA : six axis control mode • Different setting following the mission phase MLT • MCA sensors : – Star Tracker – T-SAGE MAS MGT2 • MCA actuator : – CGPS MCA MCA6 MCA3 MSP MCAc MCAi MCAs XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 14 Satellite description ■MCA control loop MCA software (4Hz sampling) Disturbing Forces & Torques F Drag free com control laws Thrusters + Fc i Cold Gas F + F Satellite selection = thrusters C C i 8:1 THRUST T dynamics ωc - Attitude logic Qc + control laws Tcom Estimates : Attitude measurement attitude Estimation Star Tracker Qˆ sat filter rˆ γ ω sat rˆ Angular Acceleration Bias_acc Measurement ω& ω& m 6-axis accelerations γ accelerometer Linear acceleration measurement m XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 15 Satellite description GDM ECM ■Cold Gas Propulsion System Tanks Composed by two identical sub-modules PRM • HP Fill&Drain Valve Propellant : 8.25 kg of N 2 each Passivation HP Filter 5µm Valve Gas Opposite HP Pressure Transducers • Thruster actuated in continuous mode (at 4 Hz) Drains HP Latch Valve • Thrust regulation by an internal control loop HP Filter 1µm Main requirements Sonic Restrictor Double Stage Pressure • Thrust range between [0:300] µN Regulator • Thrust resolution less than 0.2 µN LP Pressure Transducer • LP Test Port Thrust axial noise less than 3.22 µN rms Plenum in [0.001:10] Hz bandwidth LP Filter 5µm LP Latch • Thrust linearity less than 5% Valves • Time response of 250 ms at 1 σ TRM LP Filter TCS TCS TCS TCS TCS TCS TCS TCS Micro in total thrust performance range PMV PMV PMV PMV PMV PMV PMV PMV Thruster Developed by CNES and ESA • Design based on under development or on the shelf components and equipments (except for electronic module) XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 16 Satellite description ■Passive desorbitation system IDEAS Based Gossamer structure • Deploying 2 arms of 4.4 m long Increase the drag surface of the satellite • Geometry of deployment optimized Reduce the re-entering time • Without IDEAS : 71 years • With IDEAS : 28 years XXV th Annual AIAA/USU Conference on Small Satellite, Logan UT , 8th August 2011 17 Satellite performances ■Satellite budget Dimension (in stowed configuration) • X × Y × Z : 1.380 × 1.040 × 1.500 m MCI • Mass max : 298 Kg (i.e.